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NASA Contractor Report 3260

150 and 300 kW Lightweight Diesel Design Study

Alex P. Brouwers

Teledyne Continental Motors Muskegon, Michigan

Prepared for Lewis Research Center under Contract NAS3-20830

N/ /X National Aeronautics and Space Administration

Scientific and Technical Information Office

1980

TABLE OF CONTENTS

PageNo.

1.0 Summary ...... 1

2.0 Introduction ...... 4

2.1 Advantages of the ...... 4

2.2 Previous Aircraft Diesel Engines ...... 5

2.3 Scope of the Project ...... 7

2.4 Relative Merit of this Project to the General Field ...... 7

2.5 Significance of the Project ...... 7

3.0 Engine Design Study ...... 8

3.1 Technology Analysis ...... 8

3.1.1 Literature Search ...... 8

3.1.2 Definition of the Technology Base ...... 8

3.1.3 Definition of the Design Approaches ...... 16

3.1.4 Criteria Attributes ...... 17

3.1.5 Ranking Priorities ...... 18

3.1.6 Rating of Criteria ...... 19

3.1.7 Logic of Ranking ...... 19

3.2 Choice of EngineConfiguration and Technologies ...... 27

3.2.1 Initial Elimination of Items from the Flow Chart Figure 3-1 ...... 27

3.2.2 Choice of Engine Configuration ...... 29

3.2.3 Comparison of 2- Cycle Operation vs. 4-Stroke Cycle ...... 34

3.2.4 Final Engine Configuration ...... 35

3.2.5 Choice of Technologies ...... 36

3.3 The 298 kW 6- Engine ...... 36

3.3.1 Stroke Cycle ...... 36

3.3.2 Uncooled Cylinders ...... 36

3.3.3 Injection System ...... 36

3.3.4 Independent Operation ...... 36

3.3.5 Synthetic Oil ...... 39

iiio°° Page No. 3.3.6 Initial Performance Parameters ...... 39

3.3.7 Engine Concept Design ...... 40

3.3.8 298 kW Engine Operating Data ...... 47

3.3.9 P-V Diagrams ...... 48 313.10 Stress Calculations ...... 54

3.3.11 Projection of Fuel Consumption ...... 58

3.3.12 Energy Balance Turbocharger -- Take-Off ...... 60

3.3.13 Cooling Requirements ...... 61

3.3.14 Anticipated Maximum Surface Temperatures of Engine Components ... 63

3.3.15 Weight of the 298 kW Diesel ...... 63 3.3.16 Initial Cost of the 298 kW Diesel ...... 64 3.3.17 Emissions ...... 65

3.3.18 Noise ...... 65

3.3.19 Risk Areas Associated with the Selected Design ...... 66

3.3.20 Proposed Development Program for the 298 kW Diesel Engine ...... 66

3.3.21 Alternate Technologies ...... 69

3.3.22 Comparison of the 298 kW Aircraft Diesel and a Comparable Current Engine ...... 70

3.4 The 149 kW 4-Cylinder Engine ...... 72

3.4.1 Technologies Applied to the 149 kW Engine ...... 72

3.4.2 Minimum Cylinder Cooling ...... 72

3.4.3 Variable Piston ...... 74

3.4.4 Mechanically Driven Centrifugal Blower ...... 75

3.4.5 Glow Plug Starting Aid in Cylinders ...... 75

3.4.6 Direct Drive ...... 75

3.4.7 Initial Performance Parameters ...... 76

3.4.8 Engine Concept Design ...... 76

3.4.9 149 kW Engine Operating Data ...... 87

3.4.10 P-V Diagrams ...... 88 3.4.11 Stress Calculations ...... 93

3.4.12 Projection of Fuel Consumption ...... 97

3.4.13 Cooling Requirements ...... 97

iv Page No. 3.4.14 Anticipated Maximum Surface Temperatures of Engine Components ... 98

3.4.15 Turboc harger Operation ...... 100

3.4.16 Blower Operation ...... 100

3.4.17 Weight of the 149 kW Diesel ...... 100

3.4.18 Initial Cost of the 149 kW Diesel ...... 100

3.4.19 Emissions ...... 101

3.4.20 Risk Areas Associated with the Selected Design ...... 101

3.4.21 Proposed Development Program for the 149 kW Diesel Engine ...... 102

3.4.22 Comparison of the 149 kW Aircraft Diesel and a Comparable Current Gasoline Engine ...... 105

4.0 Engine/Airframe Integration ...... 107

4.1 Engine Installation ...... 107

4.1.1 Description of the Layouts ...... 107

4.2 Aircraft Configurations ...... 114

4.2.1 Twin Engine Airplane ...... 114

4.2.2 Single Engine Airplane ...... 116

4.3 Aircraft Performance Evaluation ...... 118

4.3.1 Program Input Data ...... 118

4.3.2 Calculation Method ...... 122

4.3.3 Results of the Simulation Program ...... 123

4.4 Operating Cost Estimates ...... 124

4.4.1 Airplane Acquisition Cost Estimates ...... 12q

4.5 Propeller Noise Estimates ...... 127

5.0 Conclusions ...... 128

6.0 Recommendations ...... 130

7.0 List of References ...... 131

Append ixes A - Bibliography ...... 132 B - Metric Conversion Factors ...... 143

v

1.0 SUMMARY

Energy conservation, uncertainties of fuel supply and limited availability of high octane gasoline, have renewed the interest in the diesel aircraft engine, since its fuel economy is better than any type of aircraft engine currently in production.

Aircraft diesel engines have been developed before, notably the "JUMO," the Napier "NOMAD" and the McCulloch TRAD 4180. Of these, only the Junkers opposed piston, 2-stroke cycle engine ever reached the production stage. The was a 2-stroke cycle, turbocompounded design. Its complexity and the fact that it invaded the territory of turbine engines probably accounted for its demise. The McCulloch engine came close to flying when the program was terminated for non-technical reasons.

New technologies, now under active development, will result in even better fuel economies than can be obtained with current state-of-the-art diesel engines. These technologies also make it possible to develop a powerplant which is more compact and lighter than current gasoline aircraft engines.

Two engines were investigated in the study, a 298 kW diesel for a twin engined airplane and a 149 kW diesel for a single engined aircraft.

The study consisted of three major phases:

1. Technology Analysis. All in-depth survey of available aviation and automotive sources was conducted to identify new developments which offer potential benefits to an aircraft engine.

The technology base includes definition of:

A. Existing automotive diesel technology, extrapolated to the expected level in the late 1980's.

B. Existing and extrapolated aircraft engine technology.

C. On-going diesel aircraft engine developments.

These technologies were then evaluated and ranked on the basis of performance and adaptability.

. Engine Concept Design. The technologies which were chosen as a result of the evaluation and ranking process were applied to the design of the 149 and 298 kW engines. Performance, stress, weight, and cost calculations were made concurrently.

. Engine/Aircraft Integration Study. The results of Phase 2 were then used in an engine-aircraft integration study to determine the performance improvement of an airplane equipped with diesel engines. The study indicates that the diesel promises to be a superior powerplant for aircraft. The following tabulation, in which the 298 kW diesel is compared to a comparable gasoline aircraft engine shows a reduction of fuel flow, a smaller package and reduced engine weight; see Table I.

TABLE I Diesel vs. Gasoline Engine

4-Cycle 2-Cycle GTSlO-520-H Diesel

Configuration 6-Cyl. opposed 6-Cyl. radial x Stroke mm 133.35 x 101.60 100 x 100 Displacement liter 8.514 4.712 Take-off Power kW 279.64 298.28 RPM at Take-off 3400 3500 Fuel Flow at Take-off kg/hr 119.07 67.13 65% Cruise Power kW 181.76 193.88 Fuel Flow at Cruise kg/hr 49.75 37.74 Dimensions: Length mm 1657 1105 Width mm 865 632 Height mm 680 660 Dry Weight kg 262 207

The superior characteristics of the diesel powerplant result in a much improved aircraft performance. The following tabulation shows the performance of a twin engine aircraft equipped with gasoline engines or diesels. Payload is increased by 8% and, simultaneously, the range is extended by 50%:

TABLE II Aircraft Performance

Gasoline Diesel Powered Powered Rated Power kW 298 298 Max. Take-off Weight kg 3671 3671 Std. Empty Weight kg 2380 2294 Useful Load kg 1291 1377 Useable Fuel kg 609 652 Payload kg 683 726 Max. Cruise Speed km/hr 454 472 Range km 1805 2592

A similar performance improvement is obtained with the 4-cylinder 149 kW diesel engine. The technologies which result in this high level of performance are, although advanced, not untried. The adiabatic engine, the catalytic combustor and the high-speed alternator are currently under development under various contracts. It should be noted here that, although the concept engine proposes the use of ceramic combustion system components, the use of such materials for "man rated" aircraft may be 20 years away. These were included primarily to show what may be ultimately possible. However, alternate solutions are given which will result in a small reduction of performance but nevertheless will result in a power plant which far out performs the gasoline aircraft engine. 2.0 INTRODUCTION

2.1 Advantages of the Diesel Engine

The current trend of ever increasing fuel prices and the dependence on imported fuel dictate the use of powerplants that offer the best in fuel economy. The diesel engine has always been burdened with the stigma of being heavy, thus offsetting its advantage of low fuel consumption for aircraft applications. If it is possible to build an engine that combines low fuel consumption and low weight, then that engine becomes a very attractive aircraft powerplant. Old and once discarded concepts can become attractive by applying new technologies.

A conventional diesel engine requires high compression ratios for starting and low load operation. This results in high firing pressures at full load when the engine could run at a much lower compression ratio. New technologies make it possible to combine good startability with low firing pressures at full load. The study shows that the weight of the diesel can be reduced below that of current gasoline aircraft engines.

The diesel engine offers more advantages in addition to low fuel consumption:

1. Lower operating cost: • Lower cost of fuel • Reduced maintenance • Extended TBO

2. Greatly reduced fire and explosion hazard.

3. Better in-flight reliability. No ignition and mixture control problems.

4. Multifuel capability. The engine will be capable of burning a variety of fuel.is, to be discussed in detail in Section 3.1.4.

5. No inlet icing problems.

6. Improved altitude performance. The 298 kW engine will be capable of continuous full power operation at 6150 m. altitude.

7. Safe cabin heating from exhaust stacks, less danger of carbon monoxide.

8. Exact fuel metering indicator. The rack position determines the fuel flow.

9. Fewer controls for the pilot: • No mixture control • No inlet heat control • No manual waste gate • No mandatory power reduction.

10. No electrical interference from ignition system.

4 2.2 Previous Aircraft Diesel Engines

Table III shows a listing and design data of aircraft diesel engines. No clear trends follow from this tabulation Seven of the thirteen engines have a radial configuration, seven were air-cooled, eight were 2.stroke cycle.

The tabulation becomes more meaningful if specific ratios are used. See Table IV.

Formulas used in the tabulations are:

2-Stroke cycle power:

BMEP x D x RPM kW = 60,000

4-Stroke cycle power:

BMEP x D x RPM kW = 120,000

BMEP is expressed in kPa

Engine displacement D in liters

s x RPM m/sec Piston speed Vp _ 30

s = stroke in meters

Some observations can be made from the Tables III and IV. Average specific weight values are:

4-Stroke cycle engines 1.408 kg/kW 2-Stroke cycle engines 1.071 kg/kW Air-cooled engines 1.277 kg/kW Liquid-cooled engines 1.082 kg/kW

The numbers indicate that a 2-stroke cycle engine can be expected to be lighter than a 4-stroke cycle engine. A comparison of air-cooled and liquid engines would seem to favor the liquid-cooled engine. However, the engine weights of liquid-cooled engines do not include the weight of the cooling package, which accounts for approximately .160 kg/kW. The corrected values then become:

Air-cooled engines 1.277 kg/kW Liquid-cooled engines 1.242 kg/kW TABLE III Previous Aircraft Diesels No. Bore Stroke Displ. Compr. Power Wgt. Make Model Config. Cycle Cooling Cyl. mm mm ,e Ratio kW RPM kg Year 1. DR980 Radial 4 air 9 122 152 16.1 16:1 174 2050 231 1930 2. Guiberson A980 Radial 4 air 9 122 152 16.1 14.7:1 155 2050 231 1931 3. Deschamps 30°A 2 liquid 12 152 229 50.5 16:1 1000 1750 1089 1934 4. Bristol Phoenix Radial 4 air 9 146 190 28.75 14:1 318 2000 494 1934 5. Zbrojovka ZOD Radial 2 air 9 120 130 13.2 15:1 207 1600 297 1935 6. Hispano Clerget 14F2 Radial 4 air 14 140 160 34.5 15:1 518 2200 600 1935 7. Salmson SH18 Radial 2 air 18 118 150 29.5 16:1 481 1700 567 1935 8. Mercedes OF2 60 ° V 4 liquid 12 165 210 53.9 15:1 592 1790 935 1935 9. Junkers 204 Opposed 2 liquid 6 120 2 x 210 28.75 17:1 570 1800 750 1935 10. Junkers 205 Opposed 2 liquid 6 105 2 x 160 16.6 16:1 444 2200 510 1936 11. Junkers (1)* 207 Turbo Opposed 2 liquid 6 105 2 x 160 16.6 16:1 740 3000 649 1938 12. Napier (2)* Nomad Flat 2 liquid 12 152.4 187.33 41.0 16:1 1984 2050 1624 1953 13. McCulloch (3)* TRAD-4180 Radial 2 air 4 98.43 98.43 3.0 15:1 150 2850 149 1970

*Numbers in parentheses refer to list of references at the end of this report.

TABLE IV Specific Data of Previous Aircraft Diesels Piston Piston Heat BMEP Speed Load Spec. Power Spec. Wgt. Make Cycle SIB kPa mlsec kWlcm z kWl_ kWlkg

Packard 4 1.246 633 10.39 .165 10.81 .75 Gu i berso n 4 1.246 564 10.39 .147 9.63 .67 Deschamps 2 1.507 679 13.36 .459 19.80 .92 Bristol 4 1.301 664 12.67 .211 11.06 .64 Zbrojovka 2 1.083 588 6.93 .203 15.68 .70 Hispano 4 1.143 819 11.73 .240 15.01 .86 Salmson 2 1.271 575 8.50 .244 16.31 .85 Mercedes 4 1.273 736 12.53 .231 10.98 .63 Junkers 204 2 1.750 661 12.60 .420 19.83 .76 Junkers 205 2 1.524 729 11.73 .427 26.75 .87 Junkers 207 2 1.524 892 16.00 .712 44.58 1.14 Napier 2 1.229 1416 12.80 .906 48.39 1.22 McCulloch 2 1.000 1053 9.35 .493 50.00 1.01

6 2.3 Scope of the Project

The purpose of the study is the conceptual design of two advanced diesel aircraft engines and the integration of these engines into airframes which are optimized for their use. One engine of 149 kW is designed to power a light single engine aircraft, the other of 298 kW is designed to power a heavy twin engine aircraft. The engines are designed to result in aircraft performance as shown :

Aircraft Characteristics Single Engine Twin Engine

Design Payload 2 passengers 3 passengers 181 kg 272 kg Max. Payload 4 passengers 6 passengers 363 kg 544 kg Design Speed kmlhr 240 400 Design Range km 1370 1610 Design Cruise AIt. m 3050 6100 Take-off to 15 m 520 1100 (standard day) Climb Requirement FAR Part 23 FAR Part 23

The results of a "GATE" computer simulation show that the aircraft will exceed these requirements by a wide margin.

2.4 Relative Merit of this Project to the General Field

Sizeable reductions in fuel consumption -- 47% at take-off and 29% at 65% cruise power are projected for the proposed 298 kW design. This results from the use of ceramics (uncooled cylinders) and a high efficiency, high pressure ratio turbocharger. These techniques when developed can be equally applied to any diesel powerplant to obtain reductions in fuel consumption from present levels.

Other technologies which are applicable to the general diesel field are: • The catalytic combustor • The high speed starter/alternator • Operation of the turbocharger independent of the engine

2.5 Significance of the Project

The importance of the program is the contribution this new engine will make in reducing the fuel consumption of general aviation aircraft. Also, the capability of the proposed engines to operate on a variety of fuels will make the diesel-engined aircraft less vulnerable to local scarcities of a particular fuel.

The significance of the program goes beyond aircraft engines since the technologies to be developed in this program would be applicable to other fields as well. 3.0 ENGINE DESIGN STUDY

3.1 Technology Analysis

3.1.1 Literature Search

An in-depth survey of available aviation and automotive sources was conducted to identify new or expected developments which offer potential benefits to a general aviation aircraft engine. A listing of this material is enclosed with this report as Appendix A.

3.1.2 Definition of the Technology Base

Following a study of the literature, a schematic, Figure 3-1, was made which shows all the technologies and the interrelationship of configurations that can be considered for an advanced diesel aircraft engine. The methodology of evaluation and ranking is presented in the definition of design approaches section.

1. Engine Configuration. Piston engines have been built in radial and in-line configurations. The has a weight advantage, the in-line engine has reduced frontal area.

2. 2-Stroke Cycle vs. 4-Stroke Cycle. The following 2-stroke cycle scavenge systems were considered, see Figure 3-2:

A. Loop Scavenging. This is a valveless configuration in which the piston controls both intake and exhaust timing events by covering and uncovering of two sets of ports located near the bottom end of the piston stroke.

B. Uniflow Scavenging. This is a system where fresh air is admitted at one end of the cylinder and exhaust gas is discharged _at the other end. No short circuiting of air flow between intake and exhaust ports is possible with this system. Three arrangments are possible for uniflow scavenging: • Opposed pistons • Twinned cylinders • Inlet ports and exhaust valves

3. Combustion Systems. Four systems were considered -- Figure 3-3.

• Direct injection.

• Prechamber, where the fuel is injected and ignited in a high turbulence system.

• MAN System. Essentially a prechamber built into the piston.

° NAHBE System. (4)* (Naval Academy Heat Balanced Engine). This system is based on a pressure exchange between an annular space in the piston and the main combustion chamber. _=.._,,.,L_==:.,, ...... I-----4:::"

, i IN-LINE L S:=:- =

r...... t!iEi: !...... ] I I RrRCR_FT D;ESEL ...... 1 i i l== _:_:_::: L...... t:,i:;t

RRDIRL

I

FIGURE 3-I AIRCRAFT DIESEL DESIGN APPROACHES \ /-i}_

PRECHAMBER

DIRECT INJECL_ON

2S.C. UNIFLOW SCAVENGE OPPOSED PISTONS

MAN 2S.C. 2S.C. SYSTEM UNIFLOW SCAVENGE UNIFLOW SCAVENGE TWINED CYLINDERS PORTS AND VALVES

FIGURE 3-3

FIGURE 3-2 SCA VENGE SYSTEMS COMBUSTION SYSTEMS

x HIGH COMPRESSION RATIO

z LOW o OMPRESSION U_ _0 RATIO

0 / I u / I STRUCTURAL / I LIMIT / ./-- / 1 / / u.J J,c A/ / /'B ,_, _/) J _=.=, J J u.,_: J jf J tO0 PERCENT LOAD

FIGURE 3-4 FIGURE 3-5

VA RIA BLE COMPRESSION RA TIO PISTON COMPRESSION RA TIO AND FIRING PRESSURE

10 4. Compression Ratio. A low compression ratio results in low firing pressures and thus a lighter engine. The disadvantage is that the compression temperature becomes too low for ignition at starting and idle operation. Several methods are available to overcome this problem:

• A flame heater in the intake manifold.

• A combustor in the exhaust manifold to accelerate the turbocharger at low engine speeds.

Variable compression ratio piston, Figure 3-4 (5)*. Regulation of the maximum pressure in the oil chamber results in a limitation of cylinder firing pressures. The firing pressure at full load is indicated by point C, Figure 3-5. Without VCR this would have been point D.

5. Configurations. A conventional crankshaft consists of main journals, crankpins, and cheeks. A barrel crankshaft has main journals and cheeks combined into large discs. This results in a shorter shaft and a higher torsional natural frequency.

6. Injection Systems:

• Conventional, consisting of injection pumps and pressure operated injectors.

• Unit injectors. Pumps and injectors are combined in one unit.

• Hydraulically amplified injectors (UFIS).

• High pressure system (CAV. (6)*

• Piezo-electric operation.

7. Degree of Cooling. Tests at TCM/GPD on air-cooled cylinders have shown that the cooling air flow can be reduced without harmful effects to the integrity of the cylinders. In aircraft engines this results in a reduction of the cooling drag.

A program is in progress to develop an engine without cylinder cooling, the "Adiabatic" diesel engine. (7)* This program is under way at Cummins Engine Co. and is sponsored by the U.S. Army Tank Automotive Research & Development Command. The cylinder and piston temperatures necessitate the use of ceramics.

*Numbers in parentheses refer to list of references at the end of this report.

11 8. Turbocharging Systems for 2-Stroke Cycle.

A° Turbocharger and Geared Blower. The 2-stroke cycle requires a positive pressure differential between intake and exhaust manifolds in order to accomplish the scavenging of the cylinders. The turbocharger, however, produces a negative pressure ratio at low engine power. A geared blower, therefore, is required for starting and low load operation. A clutch can be provided to disengage the blower once the turbocharger comes up to speed.

B° Geared Turbocharger and Clutch. In this case, the turbocharger is mechanically connected with the crankshaft. The clutch provides free shaft operation when the turbocharger is capable to provide a positive pressure ratio.

C° Geared Compressor and Geared Turbine. This system guarantees adequate air flow to the cylinders, while any excess exhaust gas energy not needed to drive the compressor becomes useful shaft energy.

D° Comprex. (8)* Figure 3-6. This is a device in which the pressure pulses of the exhaust gases are directly transferred to the induction air. The rotor is mechanically driven but no mechanical energy is required to compress the air.

J HIGH PRESSURE EXHAUST GAS

HIGH PRESSURE

LOW PRESSURE AIR

LOW PRESSURE EXHAUST GAS

TRAVEL OF AN EXHAUST GAS PARTICLE

_////'_J .m_ ROTOR I_ ,_/_ • MOTION _/

FIGURE 3-6 COMPREX

12 E. Turbocharger Operation Independent of Engine Operation. (9)* In this system, a starter motor engages with the turbocharger. An injector and combustor are located between the compressor and the turbine. Similar to a gas turbine, at some speeds the system will become self-sustaining without the aid of the starter motor. The high pressure, high temperature air from the compressor will then be utilized to start the engine. The system will be explained in more detail when the 298 kW engine is described.

Fo Differential Compound. Figure 3-7. This system compensates for a sudden increased power demand on the output shaft. This demand will slow the output shaft down and at a constant engine speed will result in a speed increase of the compressor. The resulting higher intake manifold pressure means a higher engine-torque capability to meet the increased power demand.

G. Diesel Rankine ; (10)* This system utilizes the heat energy in the exhaust gases for a secondary steam cycle. The steam is expanded in a turbine which is mechanically linked to the engine output shaft.

9. Turbocompounding. Figure 3-8. Exhaust gas first expands in the first stage free-shaft turbine which drives the compressor. The second stage turbine is mechanically connected with the engine crankshaft.

OUTPUT

ENGINE

FIGURE 3-7 DIFFERENTIAL GEARED DRIVE

13 EXHAUST FROM POWER TURBINE

/ /

\ EXHAUST TO POWER TURBINE

COMPRESSOR AiR INLET COOLING FAN CLUTCH ASSEMBLY

4_

EXHAUST TO TUI_BINE FROM PISTON ENGINE

COMPRESSED AIR TO PISTON ENGINE

,, j;

FIGURE 3-8 TURBOCOMPOUNDING 10. Catalytic Combustor. Figure 3-9. The catalytic combustor consists of a fuel injector, igniter, flame holder and a catalyst. An infrared surface heater, capable of generating sufficient heat to activate the catalyst could be incorporated. The catalytic combustor performs several functions:

• It provides the means to run the turbocharger independent of the engine.

• It provides additional thermal energy to the turbocharger turbine at maximum power operation.

• It reduces exhaust emissions.

ENGINE EXHAUST

ELEMENT

HEATER CCE) C2) AIR m_ CATALYST TO TURBOCHARGER CZD IGNITER' ' r--I-

FUEL

FIG URE 3-9 CA TA L YTIC COMBUS TOR

11. Variable Area Turbocharger (VAT). Figure 3-10.

A. Compressor Section. High pressure ratio are severely limited in flow range unless some form of flow regulation is applied to the compressor section. The problems are surge and choking in the inducer and diffuser. Flow limits can be increased if the area which is choked can be increased. Similarly, surge flow limits can be reduced if the area where surge occurs can be reduced. The variable diffuser vane concept can provide the combination of high pressure ratios, high efficiencies and a wide flow range.

g. Turbine Section. The function of the turbine is to meet the varying power demands of the compressor. In order to meet the wider range of the variable compressor diffuser, it is necessary to provide a variable area nozzle on the turbine.

15 \

FIGURE 3-10 VARIABLE AREA TURBOCHARGER

3.1.3 Definition of the Design Approaches

An evaluation and ranking procedure of the technology base was devised to arrive at a definition of the design approaches for both engines. It was not possible due to time limitations to use the NASA Systems Engineering Decision Algorithm. Instead, a simplified ranking procedure was used.

The following criteria were observed:

1. The engine must be a piston-crankshaft type powerplant.

2. Be compatible with conventionally designed aircraft (size and drag).

16 3. Allow manufactureof an experimentalmodel in five years, 4. Be readyfor productionin the late 1980's. 5. Meet EPA1979emission standards(guidereferenceonly). 6. Havemulti-fuelcapability.

7. Haveengine performancecomparableto current aircraft engine. 8. HavelowerBSFCthan presentengines.

9. Maximumspecific weightsof .852 kg/kW for the 298 kW engine and 1,095 kg/kW for the 149 kW engine.

10. Life cycle costs equal or less than present aircraft engines.

11. Avoid problem areas encountered in current aircraft engine designs.

3.1.4 Criteria Attributes

1. Performance. The engines must be capable of meeting the propulsion requirements for a given air frame application.

2. Weight, Size and Center of Gravity.

A. Weight. Engine weight must be held to a limit where aircraft performance is not adversely affected. Any extra weight relative to current engines will require redesign of engine mounts and possibly wing design.

B. Size. Those dimensions or areas which determine the contribution of the powerplant to the overall drag of the aircraft are considered most critical.

C. Center of Gravity. A large deviation of the center of gravity relative to current aircraft engines must be avoided. Such a deviation would affect the flight characteristics of the aircraft and would require a redesign of the airframe.

3. Fuel Economy. Brake specific fuel consumption will be used as a criterion of fuel economy. For this study acceptable BSFC targets at 65% cruise power will be:

A. 275 g/kW-hr for the 149 kW engine

B. 245 g/kW-hr for the 298 kW engine

17 4. Multi-Fuel Capability. The design goal of the study is efficient operation on diesel fuel and jet fuels, specifically automotive diesel fuels Dr1 and Dr2, turbine fuels JP4, JP5 and Jet A and kerosine. Operation on gasoline is possible only if an additive or lube oil is mixed with the gasoline. Engine design modifications may be developed that would allow the installation of a heater in the intake manifold. With such a device low cetane fuels may be used; however, power output and engine life would be negatively effected under the circumstances and operation in this mode should not be considered routine.

= Reliability. Reliability is defined as the probability that a subsystem or component will perform satisfactorily for the projected life of the engine or between times of inspection or overhaul.

6. Noise. Engine-related noise will be held at or below the level of current aircraft engines.

7. Technology. Technology is defined as the level of available knowledge regarding the functioning of a proposed subsystem or component.

= Life Cycle Costs. Life cycle costs include all costs related to the operation of the engine over the life of the engine, such as fuel, oil, inspections, maintenance, and overhauls.

. Component Costs. This criterion relates to the initial cost of components. It includes manufacturing costs and the launching costs, such as R&D, and tooling.

10. Integration. Integration is defined as the capability of the proposed design approach to be integrated into the overall engine design, as well as the ability to adapt the engine to conventional airframes. The engine,s cooling characteristics in the installed environment will be considered an important facet of this criterion. The center of gravity of the proposed powerplant will also be considered in this category.

3.1.5 Ranking Priorities

Decision criteria were divided into three major groups:

1. Highest priority is given to those criteria which determine that the engine is compatible with conventional airframes and that it has improved fuel economy when compared to current aircraft engines. Included in this category are: • Performance • Weight • Size • Fuel Economy

2. Next highest priority is given to those criteria that make the diesel engine a more attractive aircraft powerplant than current gasoline engines: • Multi-Fuel Capability • Reliability • Emissions • Noise • Technology

18 3. Thefollowing criteria are importantbut not to the degreeattachedto the previouscategories: • Costs • Integration • Cooling • Drag

3.1.6 Rating of Criteria

The rating system was used as a guide to identify those technologies which offer the most significant payoff, Table V.

TABLE V Weighting Factors Max. Weighting Max. Points Factor Factor x Priority -- Assigned Performance 10 10 100 Component Weight 2 10 20 Effect on Engine Weight 10 10 100 Effect on Engine Size 10 10 100 Effect on Fuel Economy 10 9 90 Engine Friction 4 2 8 Multi-Fuel Capability 8 8 64 Reliability 8 7 56 Emissions 6 6 36 Noise 6 5 30 Technology 6 4 24 Life Cycle Cost 4 3 12 Component Cost 4 2 8 Effect on Engine Cost 4 2 8 Integration 4 1 4 Cooling 4 2 8 Drag 4 1 4 Total any item 672 points

3.1.7 Logic of Ranking

Following is the reasoning that went into the assignment of points:

1. Engine Configuration.

A. Cylinder Arrangement.

a. Weight.

• In Line. Increased weight due to longer crankcase and longer crankshaft.

• 60 °, 90 °, and 120°V. Heavier crankcase than opposed cylinders due to direction of rod forces, resulting in heavier main bearing area construction.

19 • Radial Single Row. Lightest crankcase.

• Radial Double Row. Heavier crankcase than single row. b. Size. Same frontal area for Vee engines and opposed cylinders. Minimum overall size for radial single row. c. Engine Friction. Engine friction will vary with the number of crankshaft bearings. d. Technology

• In Line. No problems.

• 60 °, 90 °, and 120°V. Crankcase harder to design due to direction of forces relative to main bearings.

• Opposed Cylinders, No problems.

• Radial Single Row. No problems.

• Radial Double Row. No problems.

• Radial Twinned. Little experience to count on. e. Life Cycle Cost.

• In Line. Easy maintenance.

• Vee Engines. Less accessibility of components inside Vee.

• Opposed Cylinders. Easy maintenance,

• Radial Single Row and Twinned Cylinders. Very accessible for maintenance.

• Radial Double Row. Less accessible.

f. Engine Costs. Points assigned proportional to engine weight and complexity. g. Integration.

• 120°V and opposed cylinders come closest to current gasoline engines. 60 ° and 90°V, Probably more difficult to integrate with current airframes.

• In Line. Longer than current gasoline engines.

• Radial. Totally different configuration than current aircraft engines.

2O h. Cooling.

• Single RowRadial.Best cooling conditions.

• Single Row Twinned. Less cooling between twinned cylinders.

• Double Row Radial. Slightly less cooling of the 2nd row.

• In Line. Reduced cooling rear cylinders.

B. Air-Cooled. In-Line Configuration, Cylinders in One Block.

a. Weight. Continuous fins and one-piece mounting flange add more weight to cylinder block.

b. Reliability. Portion of combustion load is absorbed by cylinder hold- down bolts of adjacent cylinders in the case of a cylinder block.

c. Technorogy. The larger cylinder block presents more casting problems than individual cylinders.

d. Life Cycle Costs. Cheaper to replace an individual cylinder in case of scuffing.

C. Liquid Cooled. In-Line Configuration, Individual Cylinders.

a. Weight. A cylinder block allows closer spacing of the cylinders. Also, no walls separate the cylinders. Same arguments apply to size.

b. Reliability. Individual cylinders require coolant connections and thus a chance of leaks.

c. Component and engine costs lower in case of cylinder blocks due to simultaneous machining of all cylinders.

D. Propeller Drive

a. Drive off the camshaft results in low engine weight and size since it allows a high engine speed and thus a lower piston displacement. Also, the gear train is simplified. However, potential severe vibration problems in engines of this type imply an intense R&D effort.

b. Drive off the crankshaft results in a low engine speed and thus a large , reflected in the weight, size, engine cost and integration factors.

c. The geared prop drive falls between the aforementioned versions.

21 2. Power Train.

A. Type of Crankshaft. The barrel crankshaft has the main journal and the cheeks combined into a single disc. It allows much closer spacing of the cylinders and results in a high natural frequency of the crankshaft system. Consequently, engine weight, size, and cost are reduced.

g. Crankshaft Material. Advantage of nodular iron is reduced weight and manufacturing cost. Stress calculations will determine the feasibility of a cast crankshaft.

C. Crankpin Arrangement. Individual crankpins for all cylinders, required to obtain even firing of an opposed 6-cylinder engine, results in increased cylinder spacing and, therefore, increased engine weight, size and cost.

D. Connecting Rod Configuration. Master and link rods result in reduced cylinder spacing.

E. Effect of Compression Ratio. High C.R. results in heavier components and increased engine weight. BSFC is lower due to higher thermal efficiency. The requirement of VCR pistons or a flame heater system in the case of the low C.R. engine resulted in lower factors for technology and life cycle cost.

F. Piston Material. Composite material will be considered in the case of an adiabatic engine, resulting in a lighter engine and lower BSFC.

G. Connecting Rod Material. Lighter composite rods will reduce average bearing loads and increase natural frequency of the crankshaft system.

H. Cylinder Construction. Current construction of cylinders, aluminum with a steel sleeve, is probably inadequate for diesel operation. Best construction is a forged steel barrel with cast-on fins. Compromise is a cast barrel, which is less expensive to manufacture.

Piston Ring Material. Composite material to be considered in the case of an adiabatic engine. Ductile iron becomes inadequate at higher operating temperatures.

J. Gear Train Location. Advantage can be taken of the offset of opposing cylinders when gear trains are located at both ends of the engine, resulting in reduced engine length. However, more gears mean increased weight and less reliability.

3. Induction System.

A. 4-Stroke vs. 2-Stroke Cycle. The simplicity or absence of the valve train in the case of 2-stroke cycle engine results in significantly reduced weight and size and more reliability. However, more effort is required to develop optimum scavenge conditions.

22 a. 2-Stroke Cycle Systems. Loop scavenging is the simplest but also the least efficient system. Opposed piston uniflow, although very efficient has the disadvantage of complexity due to two and related gearing. Twinned cylinders offer an advantage in weight and size but disadvantages of high pumping losses between the cylinders and an extensive combustion chamber development program. Uniflow with inlet ports and exhaust valve has the disadvantage of requiring a camshaft and camshaft drive and as a result increased frontal area.

C. Types of Valve Actuation. Hydraulic valve action requires pumping elements and valve actuators. Wave propagation between pump and cylinder results in valve timing which varies with engine speed. Reliability suffers because of possible leaks of fittings. Electrical valve action has advantages of simplicity and compactness.

g. Overhead Camshaft vs. Push Rods. The overhead camshaft eliminates push rods and rocker arms. Also, the natural frequency of the valve train is higher which reduces the chance of separat-ion.

E. Camshaft Material. Nodular iron is preferred because of lower cost and reduced weight. Stress calculations will decide whether a cast camshaft can be utilized.

F. Cam Followers vs. Hydraulic Tappets. Cam followers are preferred because of lower cost and higher reliability.

4. Exhaust System.

A. Cooled vs. Hot Exhaust Port. The hot exhaust port is to be considered in particular in the case of the adiabatic engine. Much development work needs to be done if the engine is equipped with exhaust valves to provide valve cooling and to ensure proper seating of the valve.

B. Exhaust Manifold. The insulated exhaust manifold will be considered in conjunction with the adiabatic engine.

5. Cooling System.

A. Liquid Cooling vs. . Reliability of the liquid cooled engine has been rated low because of the possibility of cooling system leaks.

B. Degree of Cooling. The highest score obviously goes to the adiabatic engine. The technology factor scores low due to the amount of development work that needs to be done to make such a system feasible. Reduced Cooling: TCM has done considerable work on reduction of the cooling air flow around air-cooled cylinders. It was found that the cylinder cooling can be greatly reduced from current practice without harmful effect. Piston rings appear to be the critical item when the cooling is reduced.

23 6. Combustion System.

A. Combustion Chamber Design.

a. Open Chamber. The open chamber scores highest in most categories except multi-fuel capability and emissions.

b° Prechamber System. The prechamber system is mainly used in applications where emissions and multifuel capability are prime considerations. However, the pumping losses between prechamber and main chamber and the heat loss of the prechamber result in a higher fuel consumption.

C. MAN System. The MAN system in a broad sense could be considered to be a prechamber system, and has most of the advantages and disadvantages of the prechamber engine. An added disadvantage of the MAN system is the high heat load on the piston, which practically limits the BMEP to 1000 kPa.

d.'NAHBE System. Not enough is known about the endurance features of this engine. Obvious disadvantages of this system are high pumping losses between the main combustion chamber and the annular chamber in the piston and the lack of cooling of the upper part of the piston.

g. Methods to Obtain Low Compression Ratio at Full Load. Flame Heater System. In this case, the engine has a fixed, low compression ratio. Starting, idling and low load operation require preheating of the induction air in order to obtain the ignition temperature of the fuel at the end of the compression stroke. Development work needs to be done to ensure a very reliable heater system. The main disadvantage of this system is that the engine will die if the flame heater system fails during low load operation, with no possibility of restarting the engine without the heater.

VCR Approach. The engine starts and idles at a high compression ratio. The VCR piston gradually lowers the compression ratio as the engine load increases. Disadvantages are the heavier pistons and the resultant somewhat lower natural frequency of the crankshaft system. Failure of the VCR function may result in excessive firing pressures and the failure of a cylinder but does not affect the operation of the remaining cylinders.

7. Lubrication System.

A. Type of Oil. Synthetic oil is probably a must in the case of the adiabatic engine. The technology factor for synthetic oil has been reduced to indicate that more experience with this type of oil is required inour particular application.

g. Type of Oil Filter. The centrifugal oil filter has been in use on board of diesel powered ships for more than 50 years. The centrifugal filter prolongs the life of the oil indefinitely. The problem in the case of the aircraft engine is the development of a compact, lightweight, and very reliable unit. Such a filter does not now exist.

24 8. Injection System.

A. Systems.

a. Conventional System. Most commercial systems operate at 55,000 - 83,000 kPa line pressures. These systems are well developed and reliable. Unit injectors offer an advantage of simplicity but have a disadvantage in bulkiness of the injector.

b, Hydraulic Amplified Systems. Best known is the UFIS System. Its advantages are superior injection characteristics resulting in a very low BSFC. Disadvantages are complexity and bulkiness, which at the present state-of-the-art make the system feasible only for large engines.

C. High Pressure System. This system involves line pressures over 110,000 kPa. The high rate of injection and short duration of injection result in heat release during a period of the highest instantaneous thermal efficiency and thus in a low BSFC. Endurance testing of this system will be requfred to prove its durability.

d. Piezo-Electric System. This system will have a place in the engine if electronics are used in the control system. The system will have much of the advantages of the hydraulically amplified system but with less bulkiness and complexity. Disadvantage of this system is the state-of- the-art. Much development and endurance testing must be done to prove its validity.

B. Type of Fuel Filter.

The same arguments mentioned under oil filters apply to the fuel filter. Considering the life of a fuel filter, the arguments in favor of a centrifugal fuel filter are probably even weaker.

9. Controls.

Mechanical controls have the advantage of proven reliability and a high degree of technology.

Electronic controls have the advantage of being able to handle a much more complex input of variables and thus result in a much better engine response to varying conditions. Weight and size of this system are expected to be superior and the BSFC is expected to be lowest. Hydraulic controls are expected to be heavier and bulkier and to have the lowest deg4"ee of reliability.

10. Control of Torsional Vibrations.

Pendulum Dampers. These dampers have the advantage of minimum weight and a very high degree of reliability. The disadvantage, same as in the case of tuned damper, is that they are tuned to one order of vibration. Still, we consider the pendulum dampers superior to the rubber and viscous dampers. The rubber damper will lose its effectiveness in time due to deterioration of the rubber. Also, the tuning effect depends on the spring rate of the rubber, which is affected by ambient temperature. The same is true in the case of the viscous damper.

25 11. Basic 4-Stroke Cycle Systems.

A. Supercharging. This is any system that will increase the intake manifold pressure above atmospheric. Supercharged and aftercooled operation results in the highest density of the induction air, consequently, more fuel can be burned per liter of displacement.

B. Type of Supercharging. The free shaft system, or conventional turbocharging is compact and the technology welt in hand. Disadvantage is that the turbocharger acts as a restriction in the induction system at idle and low engine loads. This is avoided in the coupled system, where the turbocharger is mechanically coupled to the crankshaft. The coupled system, however, has the disadvantage of the weight, bulk and complexity of a high speed gear train.

The comprex, a mechanically driven, gas dynamic pressure exchanger avoids the disadvantages of the turbocharger. Early disadvantages of the comprex, such as limited speed range, low efficiency and exhaust gas recirculation have been partially overcome. By its nature, the comprex has an optimum efficiency at (_ne speed. Therefore, the comprex would be optimized for the aircraft's cruise speed. The comprex utilizes the exhaust gas energy to compress the air, therefore, the energy to drive the unit is minimal. The technology factor is low because much endurance testing needs to be done.

C. VAT and Independent Turbocharger Operation vs. Conventional Turbocharging. Independent turbocharger operation offers a concept for very high supercharging without excessive exhaust port temperatures. BMEP's up to 3450 kPa may be expected, resulting in a very small piston displacement. Disadvantages are the complexity of the system and lack of experience. The variable area turbocharger-VAT has the advantage of better matching of turbocharger and engine. The BSFC is lower and the engine torque remains high at low engine speeds. Disadvantage of the VAT is the weight and bulk of the control system.

D. Types of Combustors. The combustor provides turbine power at start, idle and low engine load conditions. Thus, the compression ratio of the basic engine can be lowered resulting in reduced mechanical loads on the engine. Advantages of the catalytic combustor are a wide range of operation, reduced emissions, simple control by modulation of the fuel and the fact that re-ignition does not require a pilot. The latter is a very important advantage in the case of aircraft operation.

E. Turbocompounding, Differential Compound and Organic Rankine Systems. Turbocompounding has the advantage of a higher overall thermal efficiency of the engine system but the disadvantage of the complexity of the high speed gear train.

Differential compounding produces higher torques at reduced engine speeds but this is not a particular advantage in the case of aircraft engines.

26 TheorganicRankinesystemhastheadvantageof extractingall availableenergy from theexhaustgases.Theresult isa 15%improvementof the BSFC. Disadvantageis thecomplexityandweightof thesystemand lackof proven reliability.

12. Basic 2.Stroke Cycle Systems.

Most of the systems have already been discussed. Condition for 2-stroke cycle operation is an intake manifold pressure which is higher than the exhaust manifold pressure throughout the engine operating range.

The turbocharger and geared blower combination is the conventional approach for high output 2-stroke cycle engines. Disadvantages are the bulkiness of the system and power required to drive the blower. The BSFC consequently is higher but the advantage is that it is a proven system.

A geared turbocharger and clutch combination overcomes the disadvantage of the turbocharger at low engine speeds .but has the disadvantage of a high speed gear train.

Another scheme consists of a geared compressor and separately geared turbine. Advantages are a positive /kP throughout the engine speed range and better utilization of the exhaust gas energy. Disadvantage is the complexity of the system.

The total points for each item were then entered on the matrix chart, Figure 3-11. The first number represents the total minus technology points, the second technology. The value given to technology represents the current state-of-the-art. Any value less than 24 indicates that some development will be required. By listing technology separately the possibility is avoided that an attractive design approach might be rejected on account of a low technology value.

3.2 Choice of Engine Configuration and Technologies.

3.2.1 Initial Elimination of Items From the Flow Chart [Figure 3-11]:

1. 2-Stroke Cycle Comprex: Comprex cannot produce boost/back pressure _P required for 2-stroke cycle operation.

2. Coupled Turbocharged Systems: Considered too complex and too heavy relative to the benefits in the case of small powerplants.

3. In=Line Configurations: Too long, too heavy.

4. 60°V Configuration: Insufficient space inside V for injection pump and intake manifolds.

5. Cylinder Block Air-Cooled In-Line Engines: Too complex, probably too heavy. High replacement costs in case of a cylinder failure.

27 6. Liquid-CooledIn-Line:Too proneto leakageproblems.Weight penaltyof radiator, water pumpand hoseconnections. 7. Uniflow Scavenging:Complexityand weightdisadvantageare not justified when comparedto valvelessloop scavenging.

8. Masterand SlaveRods:Difficult to use in combinationwith VCRpistons. Veryhigh unit pressuresbetweensecondarypins and pin boresin masterrod in caseof high diesel combustion pressures.

9. PropDriveOff Camshaft:Resultsin very highcrankshaftand camshaftstressesdue to torsional vibrations.

10.GearTrain Both Endsof Camshaft:Increasedweight and less reliabilityoffset the advantageof somereduction in enginelength. 11.VAT:Purposeof VATis to broadenefficient operationof the turbochargerovera wide rangeof enginespeeds,not requiredin aircraft operation.A variableturbine nozzlemay.berequiredto maintiana positiveboost/B.P./k p throughout the engine load range in 2-stroke cycle operation.

12. Geared Compressor and Geared Turbine: Rejected because of complexity, weight penalty, and unreliability of 2-high speed gear trains.

13. Conventional Combustor: Rejected because of higher level of emissions compared to catalytic combustor.

14. MAN System: Developed for multi-fuel operation. High heat load on piston limits BMEP.

15. NAHBE System: High pumping losses and lack of cooling of the upper part of the piston.

16. Aluminum Cylinder and Cast-in Steel Sleeve: Structurally inadequate for the higher firing pressures of a diesel engine. Some new aluminum alloys will be explored in detailed study. • Gasoline engine 4100-5500 kPa • Diesel engine 8300-9650 kPa

17. Conventional Cylinder Cooling: The use of improved materials permits a large reduction of the cooling air flow with a resultant reduction of the cooling drag.

18. Twinned Cylinders: High flow losses can be expected in the narrow internal passage between the cylinders. Any air-swirl will be lost during passage from one cylinder to the twin cylinder with an adverse effect on combustion.

19. Centrifugal Filters: A compact, lightweight and very reliable unit will be required for aircraft use. Such a filter does not now exist.

20. Connecting Rod: The technique of producing lightweight composite rods is available and should be applied to the aircraft engine rather than steel forgings.

28 21. HydraulicValveActivation: Hydraulicvalveaction hasdisadvantagesof complexity (pumpingelementsandactuators)and reliability w possibility of leaks.Valvetiming will haveto varywith enginespeedwhichwould requirea complexcontrol system.

22. MineralOil: Expectedhigh oil temperatureswill result in a veryshort period betweenoil changes. 23. NaturallyAspiratedSystem:Low BMEPwould result in a largepiston displacement.

24. Supercharged,without Aftercooling:Aftercooling will be requiredwhenoperatingat the higher pressureratios. 25. Unit Injectors:(Pumpand Injector as OneUnit) Rejectedbecauseof weightand size.The resultantlargefrontal areawould addto dragloss. 26. PiezoelectricInjection: Rejectedbecauseof size,cost, reliability, and only marginal advantages. 27. Controls:Electroniccontrols will be appliedbecauseof the complexinput of variablesthat can be handled,resulting in a much betterengine responseto varying operatingconditions. 28. TunedDamper:Rejectedbecauseof the temperatureeffect on the rubber(alsomay not be requiredfor radial engine). Figure3-12is the resultantsimplified matrix.Thischartstill representstoo many possibilities.

3.2.2 Choice of Engine Configuration

Figure 3-13 shows all possibilities of opposed and V engines, 2-stroke and 4-stroke cycle. Shown are:

1. Firing orders.

2. Firing intervals, which indicate smoothness of operation.

3. Number of main bearings, which has a large effect on engine weight. Several configurations can be eliminated because of excessive torque fluctuations and too many main bearings for a given number of cylinders.

Tables Vl and VII show the first step in the elimination process of cylinder configurations.

The remaining candidates of Tables Vl and VII are shown in Table VIII and were checked for unbalanced piston and rod inertia forces. Only two possible in-line candidates remain: • 135 ° V8 2-stroke cycle • 90 ° V8 4-stroke cycle

Table IX lists the candidates for radial engines. The rotating and primary inertia can easily be balanced by counterweights on the crank cheeks opposite the crankpin.

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FIGURE 3-12 AIRCRAFT DIESEL DESIGN APPROACHES-SIMPLIFIED MATRIX AIRCRA -T DIESEL IN-LINE CYL NDEI AND CRANKSHAFT CONTIGURATIONS

$ 6 CVLINI)ER OPPOSED $ CYLINDER |_.oOv CVUNDER O0*V 4 CYLINDER 120*V 4 CYLINDER 900V 4 CYLINDER 90* V A CYLINDER OPPOSED 2 STROKE CYCLE ?. STROtG. CYCLE 2 STROI(£ CYCLE 2 STROt_ CVCLE 2 STROKE CYCLE Z STROVE CVCLE 2 STROKE C_/CLE

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FIGURE 3-13 AIRCRAFT DIESEL CYLINDER CONFIGURATION TABLE VI Selection of 2-Stroke Cylinder Configurations -- Step I In-Line and V Engine No.of MainBearings Objections

Flat 4 4 Large torque fluctuations. 120° V4 3 90 ° V4 3or4 Flat 6 7 Large torque fluctuations. 120° V6 7 Too many mains -- heavy crankcase. 90 ° V6 4 Flat 8 7 Too many mains. 135 ° V8 5 120° V8 5 Stepped crankpins compared to 135 ° V8. 90 ° V8 9 Too many mains.

TABLE VII Selection of 4-Stroke Cylinder Configurations -- Step I In-Line and V Engine No. of Main Bearings Objections

Flat 4 4 2 Torque fluctuations per rev. Can resonate with 2-bladed prop. 120 o V4 5 Too many mains. 90 ° V4 5 Too many mains. Flat 6 7 Too many mains. 135 ° V6 4 Stepped crankpins compared to 120 ° V6. 120 ° V6 4 90 ° V6 4 Stepped crankpins. Flat 8 9 Too many mains. 120 ° V8 5 90 ° V8 5

TABLE VIII Selection of Cylinder Configurations -- Step II In-Line and V Engine

INTERTIA FORCES AND MOMENTS FORCES MOMENTS Rot. Prim. Sec, Rot. Prim. Sec. Objections 2-Stroke Cycle 120 ° V4 0 0 :_0 :tO ;_0 ;_0 secondary unbalance 90 ° V4 0 0 :_0 0 0 ;_0 secondary unbalance 90 ° V6 0 _0 0 ;e0 ;e0 ;_0 primary unbalance 135 ° V8 0 0 0 ;e0 ;_0 0

4-Stroke Cycle

120 ° V6 0 :_0 dO ;_0 _=0 0 secondary unbalance 120° V8 0 0 ;_0 ;_0 :tO 0 secondary unbalance 90 ° V8 0 0 0 ;_0 #0 0

33 TABLE IX Radial Cylinder Configurations

RULE: 2-Stroke Cycle: Any number of cylinders per row. 4-Stroke Cycle: Uneven number of cylinders per row.

INERTIA FORCES: Rotating: Constant, in direction of crank radius. Primary: Constant, in direction of crank radius. Secondary: Zero resultant force.

Maximum number of cylinders per row is 4 due to limited bearing area of slipper connecting rods.

Candidates: 3-Cylinder single row 4-stroke cycle 6-Cylinder double, row 4-stroke cycle 4-Cylinder single row 2-stroke cycle 6-Cylinder double row 2-stroke cycle

3.2.3 Comparison of 2.Stroke Cycle Operation vs. 4.Stroke Cycle

Loop scavenge 2-stroke cycle operation offers the following advantages over 4-stroke cycle:

1. Higher specific power (kW/_)

2. Weight reduction due the absence of the valve mechanism, camshaft and camshaft drive.

3. Reduction of frontal area due to the elimination of the overhead valve mechanism.

4. Improved reliability due to fewer parts.

5. Higher engine speeds are possible due to the absence of a valve mechanism.

6. Less engine torque variation.

7. To be discussed in a later section (3.3.2), the 298 kW version is proposed to have uncooled cylinders. Valve problems in an uncooled cylinder will probably be unsurmountable; therefore, elimination of the valves is essential.

Disadvantages of 2-stroke cycle are:

1. Scavenging of the cylinders requires a positive pressure difference between intake and exhaust manifolds under all operating conditions. This means the addition of a mechanical blower in the case of conventional turbocharged 2-stroke cycle engines.

34 2. 30% moreairflow requiredfor scavengingof the cylinders. 3. Higher meancycle temperaturewhich results in a higherthermal loadingof cylindersand piston rings andraisesand NOxemissions.

4. Higheroil consumptiondueto loss of oil throughthe exhaustports. 5. Moredevelopmenttime requiredto optimizethe scavenging.

6. Coolingof the exhaustports is required. Thechoiceis the2-strokecyclesystem: 1. Theadvantageslisted aboveareof critical importancein thespecificcaseof an aircraftengine: A. Weightreduction B. Reducedfrontal drag

C. Improvedreliability 2. Thebestknowndieselaircraftenginesthat wereeverproducedor fully developed were2-strokecycleengines:

A. TheGermanopposedpiston, uniflow scavengedJunkers"JUMO."

B. TheBritish loop scavengedNapier"NOMAD."

C. TheMcCullochloop scavengedradial TRAD4180.

3.2.4Final EngineConfigurations

Thefollowing candidatesare left (all 2-strokecycle): 1. 135oV8

2. 4-Cylindersingle row radial

3. 6-Cylinderdoublerow radial This leavesonly one in-linecandidateto coverboth the 149kW and298kW.This configurationis alright for the 298kW versionbut too complexfor the 149kW engine. Our final choice,therefore,is the radial configurationasfollows:

1. A 149kW4-cylindersingle row -- 2 strokecycle. 2. A 298kW6-cylinderdouble row -- 2 strokecycle.

35 3.2.5 Choice of Technologies

The technologies now follow from Figure 3-12 taking the high score items along the "common to all versions" and "radial 2-stroke cycle" lines.

Common to All Radial 2-Stroke Versions Line Cycle Line Open chamber Individual cylinders Ceramic pistons Low compression ratio Insulated exhaust manifolds Geared prop drive No cylinder cooling Loop scavenge Tool steel piston rings Independent turbo loop Composite connecting rods Catalytic combustor Synthetic lube oil High pressure Electronic controls Conventional oil filter Conventional fuel filter Pendulum damper

3.3 The 298 kW 6-Cylinder Engine

Figure 3-14 shows an artist rendering of the proposed engine. Figure 3-15 shows the schematic of the engine. The design incorporates all the technologies which were defined before.

3.3.1 2.Stroke Cycle

The chosen system is Curtis loop scavenging. The intake ports and intake manifold are located at the propeller side of the engine, exhaust ports and exhaust manifold at the back end.

3.3.2 Uncooled Cylinders

The cylinder liner and piston top are ceramics. Tool steel piston rings will be required. Cooling air will be required only for the aftercooler, oil cooler and the cylinder fuel injectors. The exhaust ports will be oil cooled.

3.3.3 Injection System

Each cylinder receives fuel from a separate injection pump located in front of the cylinder (cool side of the engine). Failure of one pump still leaves 5/6 of engine power available. A high injection line pressure will be required to limit injection duration at high engine speeds.

3.3.4 Independent Turbocharger Operation

The turbocharger can run independent of the engine. For that purpose a high-speed starter/alternator and an oil pump are mounted on the turbocharger. A 2-way valve is placed in the intake manifold. To start the engine this valve is in the vertical position of

36 ",4

\

FIGURE 3-14 298 KW DIESEL AIRCRAFT ENGINE INJECTION PUMPS AND NOZZLES HEATER

n

TURBINE

CATALYTIC STARTER/ FUEL COMBUSTOR ALTERNATOR PUMP OIL PUMP

I VALVE I ENGINE I BLEED • ! _ COMPRESSOR

¢O I AIR STARTER EMERGENCY AIR SHUTOFF METERED ALTERNATE LUBEPUMP _ ,_ AIR SOURCE L..J TO AVOID PROPELLER ENGINE RAM ICE GOVERNOR OIL PUMP INLET AIR FILTER

• TURBOCHARGER OIL PUMP CIRCULATES OIL THROUGH ENGINE FOR PREHEATING

• BLEED AIR TURBINE CRANKS MAIN ENGINE

FIGURE 3-15 SCHEMA TIC 2-STROKE ENGINE WITH INDEPENDENT TURBO LOOP/HOT ENGINE CONFIGURA TION--NO COOLING/LO W COMPRESSION RA TIO PISTONS the schematic, which results in a turbocharger loop independent of the engine. Combustor fuel is ignited by the heater. This heater can be turned off as soon as the catalyst becomes sufficiently hot. The cycle will become self-sustaining at approximately 1/3 of maximum turbo speed, and the starter now runs as an alternator. Hot, high pressure air will flow to the engine when the 2-way valve is partially opened. The cylinder intake ports are opened during approximately 120 crank-degrees, so hot air can flow through two cylinders for preheating on an extreme cold day. The high pressure air will next be admitted to the engine-mounted bleed air starter to crank the engine. The whole sequence would be automatic on a production engine.

This system offers many advantages:

. The availability of hot induction air at start reduces the need for a high compression ratio. The engine will start and idle at a 10:1 compression ratio provided this hot high pressure air is available to it. Thus, with this low compression ratio, the firing pressures are held down to 9650 kPa at full load resulting in low engine weight.

. The engine will start easily under extreme cold conditions, a problem with current gasoline engines.

3. Hot start problems are eliminated.

4. Easy restart is available while airborne.

. The engine can be shut-off and the turbocharger kept running when the aircraft is on the ground for some period. Meanwhile, electric power, cabin heat or air conditioning remain available. This in effect converts the turbocharger into an APU.

6. The battery requirement is greatly reduced since engine cranking is accomplished by air pressure.

7. Activation of the heater greatly reduces the turbocharger response time to engine power changes.

3.3.5 Synthetic Oil

The use of synthetic oil is required due to the hot cylinders. Over the long term perhaps a method can be found to generate an airfilm between pistons and cylinder walls, in effect, using air bearing technology. This would also reduce the expected high oil consumption which is inherent to 2-stroke cycle engines.

3.3.6 Initial Performance Parameters

Table X shows the full power performance of three aircraft diesel engines. For comparison a BMEP of approximately 1100 kPa and a stroke/bore ratio = 1 were chosen for the study engine.

The engine characteristics become:

Number of cylinders 6 Take-off power 298 kW Engine speed at T.O. 3500 rpm

39 BMEP 1085kPa Displacement 4.71liter Cylinderbore 100mm Stroke 100mm Piston speed 11.67m/sec. Propellerdrive geared

Thestudy is aimed moreat the best possiblefuel economyand reliability than the highestpossible poweroutput.

TABLE X Performance of 2-Stroke Cycle Aircraft Diesel Engines Total Engine Piston Piston Spec. Spec. Spec. NO. Bore Stroke Ratio Displ. Speed Power BMEP Speed Area Power Power Weight Wt. BSFC Cyl. mm mm SIB _ RPM kW kPa mlsec. CM z kWlcm z kWlJ_ kg kWlkg glkW-hr.

McCulloch 4 98.425 98.425 1.000 3.00 2850 150 1052.6 9.35 304.3 .493 50.00 149 1.001 243 Napier-Nomad 12 152.4 187.325 1.229 41.00 2050 1984 1416.3 12.80 2189.0 .906 48.40 1624 1.222 213 Junkers-Jumo 6 105.0 2 x 160 1.524 16.62 3000 740 890.5 16.00 1039.1 .712 44.52 649 1.140 213

Average BMEP 1119.8 kPa (162.4 psi) Average Piston Speed 12.72 m/sec. (2504 fpm)

3.3.7 Engine Concept Design

The engine concept is shown in the Figures 3-16 through 3-20. The cylinders are arranged in two offset banks of three cylinders each, acting on a single crankpin. The rotating and reciprocating inertias are 100% balanced by counterweights on the crank cheeks. The pendulum dampers are mounted to the counterweights and will be tuned for the 4-1/2 and 6th orders. The cylinders are uncooled and provided with ceramic liners. The intake ports and the intake manifold are located at the front side--the cool side of the engine. The exhaust ports and exhaust manifolds are located at the backside--the hot side of the engine. Two exhaust manifolds are required to prevent the exhaust pulse of one cylinder from interfering with the scavenging of the previous cylinder in the firing sequence. The piston tops are ceramic.

The small ends of the connecting rods are designed to allow free rotation of the pistons.This should reduce the wear rate of the piston rings. The big ends of the connecting rods are designed as a slipper, i.e., each rod contacts only 1/3 of the circumference of the crankpin. This is possible for 2-stroke cycle engines because the combined load of gas pressure and inertia is always directed toward'the crankpin. The bearing material will initially be conventional, but a study could be conducted later of self-lubricating and gas bearings to eliminate the need for oil in the crankcase. The oil to be used initially is a synthetic oil which can take higher temperatures and requires fewer changes than conventional petroleum based oils.

q0 CATALYTIC COMBUSTOR ACCESSORY HOUSING / AIR STARTER

INJECTION PUMPS OIL PUMP

EXHAUST / MANIFOLD AFTERCOOLER

FIGURE 3-16 298 kW AIRCRAFT DIESEL--LONGITUDINAL SECTION INJECTORS

\

/

!

4:::

INTAKE MANIFOLD INTAKE,. • PORT__-_<.

\ / _,../ \ _EXHAUST

_ SECTION'_ THRU PORT INTAKE AND EXHAUST PORTS

FIGURE 3-17 298 kW AIRCRA FT DIESEL--CROSS SECTION 755.65 (29.75) 631.95 (24.88) FORS

GOVERNOR \

, L \ '1 \\ ! ;

L __J \ \ X AFTERCOOLER OIL COOLER

FIG URE 3-18 298 k W A IRCRA F T DIESEL-- FRON T VIE W EXHAUST MANIFOLD CATALYTIC COMBUSTOR 1104.9 (43.5) :)15 (5120) VACUUM PUMP AIR STARTER ! WITH SLIP CLUTCH INJECTION PUMPS

\

\ TURBOCHARGER ..... I SUPPORT / I

FUEL PUMP'

INDUCTION AIR FROM AIR CLEANER

INTAKE MANIFOLD ENGINE OIL PUMP

FIGURE 3-19 298 kW AIRCRAFT DIESEL--SIDE VIEW AIR STARTER WITH SLIP CLUTCH

MOUNTING HOLES ,' '\_i

\ / VACUUM PUMP

PUMP (ENGINE)

PUMP (TURBO)

{ u1 _ I

;TARTER AFTER COOLER GENERATOR

ik,..OI L COOLER ...... ,___J c -[;]J--- l

FIG URE 3-20 298k W A IRCRA F T DIESEL-- REA R VIE W Immediately in front of the 1st main bearing are 6 individual injection pumps, operated by a single lobed cam ring. Individual pumps were chosen to improve engine reliability -- failure of one pump still leaves 5 cylinders operable. Also, all fuel lines can have the same length resulting in the same injection timing for all cylinders.

A bevel gear in front of the cam ring drives the prop governor and the fuel priming pump.

A gear reduction reduces the crankshaft speed of 3500 rpm at take-off down to 2300 rpm propeller speed.

At the back of the crankcase is an accessory housing which contains the gearing for the engine oil pump, the vacuum pump, and the bleed air starter. The air starter drive is provided with a slip clutch to prevent engine damage in the case of a hydrostatic lock in one of the cylinders (accumulation of fuel due to the leakage of a fuel injector). Four engine mounting points are provided on the accessory housing. Above the accessory case is the catalytic combustor assembly. Leading to it are the two exhaust manifolds and the air bypass for operation in the APU mode.

The turbocharger is located behind the accessory housing. Figure 3-20 shows the turbine to the left and the compressor in the center. To the right is a gear housing with the high speed alternator and turbo oil pump drives.

The aftercooler and oil cooler are located below the engine accessories.

The engine will operate with a .

Speeds of accessories are shown in Table XI.

TABLE Xl Accessory Speeds

Engine 3500 rpm Propeller 2300 rpm Governor 3500 rpm Fuel Transfer Pump 3500 rpm Engine Oil Pump 3500 rpm Vacuum Pump 3500-4200 rpm Air Starter max. 10000 rpm (40 to 1 reduction) Tachometer 1750 or 3500 rpm Alternator 35000 rpm (At 70,000 rpm turbo speed) Turbo Oil Pump 8000 rpm

46 3.3.8 298 kW Engine Operating Data (11)* (12)*

Operating parameters have been calculated for the 298 kW engine and are shown in Table XlI.

TABLE XII Operating Parameters -- 298 kW Engine

100% Power 65% Power Take-off Cruise Cruise Altitude 0 6,096 6,096 meters Power 298 298 194 kW RPM 3,500 3,500 2,675 Displacement 4.71 4.71 4.71 liters Bore x Stroke 100 x 100 100 x 100 100 x 100 mm BMEP 1,085 1,085 923 kPa Compressor Pressure Ratio 4.06:1 8.30:1 6.25:1 Nominal Compression Ratio 13.185:1 13.185:1 13.185:1 ,Effective Compression Ratio 10.0:1 10.0:1 10.0:1 Barometric Pressure 101.4 46.4 46.4 kPa Ambient Temperature 15.5 - 25 - 25 °C Intake Manifold Pressure 402.4 370.2 277.6 kPa Intake Manifold temperature 116 116 116 °C Exhaust Manifold Pressure 309.5 284.8 245.5 kPa Scavenge System Curtis Loop Curtis Loop Curtis Loop Scavenge Ratio 1.3 1.3 1.3 Ratio Boost/Back Pressure 1.3 1.3 1.131 Height Intake Ports 20.65 20.65 20.65 mm Height Exhaust Ports 26.14 26.14 26.14 mm Intake Ports Open/Close 61 °47' 61 °47' 61 °47' BBDC/ABDC Exhaust Ports Open/Close 69°39 ' 69°39 ' 69039 ' BBDC/ABDC BSFC-engine 206.8 212.9 194.6 g/kW-hr. BSFC-combustor 18.2 6.1 0 g/kW-h r. BSFC-powerpack 225.0 219.0 194.6 g/kW-hr. Fuel Flow Powerpack 67.1 65.3 37.8 kg/hr Air Density .00279 .00256 .00205 kg/,_ Air/Fuel Ratio 27.50 24.59 25.47

q7 3.3.9 P-V Diagrams

Air cycle performance data has been calculated for the proposed engine. Figure 3-21 illustrates the points calculated on the P-V diagram. Specific data points for three operating conditions are given in Table XlII.

TABLE Xlll Air-Cycle Performance

100% Power 65% Power Take-off Cruise Cruise PI 356 328 262 kPa V_ .645 .645 .645 liter TI 171 171 171 °C P2 8,540 7,850 6,270 kPa V2 .064 .064 .064 liter T2 792 792 792 °C P3 9,650 8,970 7,390 kPa V3 .064 .064 .064 liter T3 932 888 982 °C P4 9,650 8,970 7,390 kPa V4 .110 .115 .111 liter T4 1,783 1,900 1,882 ° C Ps 970 960 750 kPa Vs .645 .645 .645 liter Ts 937 1023 997 °C Fuel/Cyl./Rev. .0000490 .0000504 .0000390 kg Air in Cylinder .00135 .00124 .00099 kg Q/Cyl./Rev. .502 .516 .400 kcal Qj .053 .053 .053 kcal Q2 .450 .463 .347 kcal IMEP 1,339 1,357 1,069 kPa Mech. Eff. (engine) 81 80 86 % Turbine Pressure Ratio 3.052 6.137 5.290 Compressor Pressure Ratio 4.066 8.300 6.250 Compressor Efficiency 78 79 80 % Turbine Efficiency 77 77 79 % Mechanical Efficiency 98 98 98 % Overall Turbo Efficiency 59.6 60.0 62.3 % Required TIT 616 594 458 °C Available TIT from Engine 533 569 572 °C

FLOWS: Weight Pure Air .472 .434 .266 kg/sec Weight Fuel .017 .018 .010 kg/sec Weight Exhaust Gas .489 .452 .276 kg/sec Weight Scavenge Air .142 .130 .080 kg/sec

The Figures 3-22 thru 3-24 show the schematics of these three operating conditions.

Figure 3-25 shows the engine performance curves.

48 Q2

COMPRESSION pV 1.38 = C

EXPANSION pV 1.30 = C

POINT 1 IS EXHAUST PORT CLOSING

.

1 I I I

-_Vc_ Ve r_1I

V c = CLEARANCE VOLUME

Ve = EFFECTIVE STROKE

FIGURE 3-21 ENGINE INDICATOR DIAGRAM

49 .614 kg/sec

INTERCOOLER COMPRESSOR AIR CLEANER 721 kCAL/MIN

100.2 kPa 101.4 kPa 402.4 kPa 407.4 kPa Prc = 4.066 15.5°C 115.6°C 196.1°C 15.5°C

i 101.4 kPa 309.5 kPa _ 309.5 kPa Prt = 3.052 533°C 616°C

CATALYTIC COMBUSTOR TURBINE

,631 kg / sec 298 kW 3500 RPM SEA LEVEL

FIG URE 3-22 OPERA TING SCHEMA TIC-- TA KE-OFF .568kg/sec ,q

INTERCOOLER COMPRESSOR AIR CLEANER 959 kcal/min

370.2 kPa 375.2 kPa 45.2 kPa 46.4 kPa Prc = 8.300 115.6°C 232.2°C -25°C -25°C

f_

/ u1 ,...t

284.8 kPa 284.8 kPa 46.4 kPa Prt =6.137 569.4°C 594.4°C

CATALYTIC COMBUSTOR TURBINE

v .586 kg/sec 298 kW 3500RPM 6,096m

FIGURE 3-23 OPERA TING SCHEMA TIC--IO0% CRUISE PO WER .364 kg/sec

INTERCOOLER 1400 BTU/min COMPRESSOR AIR CLEANER

277.6 kPa 282.6 kPa 45.2 kPa Prc = 6.250 115.6°C 185.6°C -25°C 46.4-25=CkPa

245.5 kPa 46.4 kPa Prt = 5.290 572.2°C

TURBINE

.356 kg/sec 194 kW 2675 RPM 6,096m

FIGURE 3-24 OPERA TING SCHEMA TIC_65% CRUISE PO WER 300

,28O 88O ENGINETORQUEAT CRANKSHAFTSPEED ,E Z 260 LLJ 840

I"" 240 800

220 FULLLOADPOWER

200

eL." I.t.J

180 a.

PROPLOADPOWER ]60

INCLUDES 140

120

BSFC 100

I I I I 2200 2400 2600 2800 3000 3200 3400 3600 CRANKSHAFTRPM

FIGURE 3-25 SEA LEVEL PERFORMANCE 6-CYLINDER RADIAL

53 3.3.10 Stress Calculations

All calculations were based on a 9650 kPa firing pressure. This pressure occurs only for short periods during take-off. Most fatigue cycles occur during cruise operation when the firing pressures and, therefore, the stresses are much lower. This results in an extra safety factor.

Figure 3-26 shows the cylinder configuration. The cylinders are arranged in two rows of cylinders. The offset of the two rows is determined by the width of a connecting rod.

The firing order is 1, 4, 2, 5, 3, 6 with even 60 ° firing intervals.

1 r-q Fq

4 6 f

i ,S ,J J 2 3 LU 5 5

FIRUGE 3-26 CYLINDER CONFIGURA TION

1. Power Train Data Weight Piston Assembly: Ceramic top .35 kg Hold-down bolt .25 kg Aluminum piston .88 kg Piston rings .12 kg Total Piston 1.60 kg

Composite connecting rod .20 kg Slipper rings .44 kg Counterweights 9.28 kg

Total reciprocating WR 24.9 kg-cm Total rotating WR 25.0 kg-cm Total counterweight WR 49.5 kg-cm

Balance 100%

5q 2. Crankshaft Stresses

A. Crankpin Fillet Radius: Max. principal bending stress 559 MPa Max. principal shear stress 120 MPa

B. Main Bearing Fillet Radius: Max. principal bending stress 87 MPa Max. principal shear stress 60 M Pa

C. Material: AMS 6415 Ultimate tensile strength (min) 1034 M Pa Endurance strength (machined & peened) 552 MPa

3. Connecting Rod Stresses and Bearing Pressures

A. Connecting Rod Max. compressive stress 291 MPa Min. compressive stress 35 M Pa Graphite-epoxy fatigue strength 390 M Pa Crankpin bearing unit load 40 M Pa SAE-794 leaded bronze max. unit load 69 M Pa Piston ball joint (30 mm_) unit load 95 M Pa Note: (w/o oil groove on ball)

B. Main Bearing (65 mm¢ x 30 mm length) Peak unit load 22 MPa Min. unit load 16 MPa

The Figures 3-27 and 3-28 show the main bearing load diagram and the crankshaft and connecting rod stresses.

4. Cylinder Barrel Stresses

Cylinder wall hoop stress 63 MPa Cylinder wall longitudinal stress 32 M Pa 8-Cylinder hold down studs M10X1.5 -- 6g Grade 8 (Proof load 40,430 N/stud) Torque to 75% proof load: 30,320 N/stud Peak dynamic load: 1,490 N/stud

Material Sintered Alpha SiC Steel

Flexural strength 442 MPa 455 MPa at 1000 oC

Density 3.16 g/cc 7.86 g/cc

Thermal Expansion 4.02 x 10-_/°C 11.4 x 10-6/°C Coefficient RT-700°C

Thermal conductivity .045 kcal/m-hr-°C 37.2 kcal/m-hr-°C at 600 ° C

55 210" 190" 170" 160" |50" 150" 170" 180" 190 ° 200" 210"

140" 220"

130' 230'

120' 240'

110' 250'

100' 260'

90" 270 =

80" 280"

140

145

50 • 310"

40" 320 ° 40 ° 320 °

330 ° 340 ° 350 ° 0 o lO ° 20" 300 30" 20= lO" 350 ° 340° 330"

FIGURE 3-27 MAIN BEARING LOAD 298kW AIRCRAFT DIESEL

56 1000 MACHINEDANDPEENED

FORGEDANDPEENED

800 MACHINED

ASFORGED

600"

CRANKPINFILLET

! -200 400 600 800 1000

MEANSTRESS-MPa

CONNECTINGROD

FIGURE 3-28 CRANKSHAFT AND CONNECTING ROD STRESS

57 5. Natural Frequencies Crankshaft System

First mode's natural frequency -- 220 Hz

1st Order 13,200 rpm 3rd Order 4,400 rpm 4-1/2 Order 2,930 rpm 6th Order 2,200 rpm

Pendulum dampers to be tuned for 4-1/2 and 6 Orders.

6. Propeller Drive Gear Stresses

A. Driven Gear 38T/7P (20 ° P.A.) 63.5 mm face width

B. Drive Gear 25T/7P (20 ° P.A.) 63.5 mm face width Wear stress 1,400 MPa Bending stress 346 MPa

Note: Overload factor taken as 1.0 Power transmitted -- 304 kW at 3500 rpm

C. Material AMS 6260 Case carburized and hardened (Rc60) AGMA allowable bending stress 414 MPa AGMA allowable wear stress 1465 MPa

3.3.11 Projection of Fuel Consumption

A comparison is made with the TCM/GPD AVCR/VAT 1360 high output 4-stroke cycle air-cooled diesel engine. This engine delivers 1120 kW at 2600 RPM. The engine was chosen because its BMEP of 2317 kPa is approximately twiCe that of the aircraft diesel, which is 1085 kPa. (A 4-stroke cycle engine of the same displacement and speed has to have double the BMEP of a 2-stroke cycle engine in order to deliver the same power.)

1. Measured AVCR-1360 performance data:

BSFC = .252 kg/kW-hr at 2600 RPM. Fuel flow 282 kg/hr (heating value 10,250 kcal/kg). Energy input 48,230 kcal/min. Heat equivalent of 1120 kW is 16,030 kcal/min or 33.2% of total energy.

AVCR-1360 cooling losses:

Cylinders 6,300 kcal/min Oil Coolers 2,140 kcal/min Aftercoolers 4,940 kcal/min Total 13,380 kcal/min = 27.7%

58 Exhaust energy loss 17,410 kcal/min = 36.1% Radiation 1,410kcal/min = 3.0%

2. Projection of an AVCR-1360 with uncooled cylinders:

The absence of cylinder cooling changes the energy balance by 6,300 kcal/min. Approximately 55% of it or 3,465 kcal/min, can be recovered as usable energy. The rest, 2,835 kcal/min, goes out the tailpipe. The new energy balance becomes:

Engine power 16,030 + 3,465 = 19,495 kcal/min. = 40.4% or 1360 kW Cooling loss 2,140 + 4,940 = 7,080 kcal/min. = 14.6% Exhaust loss 17,410 + 2,835 = 20,245 kcal/min. = 42.0% Radiation 1,410 kcal/min. = 3.0% Total 48,230 kcal/min.

Fuel flow is unchanged at 282 kg/hr

New BSFC 282 = .207 kg/kW-hr 1360

Improvement factor .207 = .821 .252

Figure 3-29 shows the energy distribution with conventional cooling and a simulated AVCR-1360 adiabatic engine.

_LOSS \ I

CONVENTIONAL UNCOOLED CYLINDERS AIR COOLING SIMULATED AVCR-1360 AVCR-1360

FIGURE 3-29 HEA T BALANCE COMPARISON

59 3. Compare to a conventional turbocharged 2-stroke cycle 8V-92T Detroit Deisel Allison engine:

Minimum BSFC = .229 kg/kW-hr (published data) At rated power BSFC = .231 kg/kW-hr (published data) Applying the BSFC improvement factor yields .821 x .229 = .188 kg/kW-hr

° This probably represents an overly optimistic number for a 2-stroke cycle engine. Therefore, a more conservative 15% improvement in BSFC over conventionally cooled engines or .85 x 229 = 194.6g/kW-hr at 65% power is projected.

5. Estimate of BSFC at take-off:

Maximum power BSFC increase over minimum fuel consumption for several

engines: Z_BSFC g/kW-hr 4-Stroke cycle VAT 1360 18.25 (data) 4-Stroke cycle Cummins LCR-V-903 12.17 (projection) 4-Stroke VAT "1790 low C.R. 12.17 (data) 2-Stroke Cycle GM 8V 92T 6.08 (data) 2-Stroke Cycle Napier Nomad 6.08 (data)

This led to the projection of a 206.8 g/kW-hr BSFC at take-off power.

The BSFC at 100% power cruise condition is expected to be higher than at take-off, 212.9 g/kW-hr, due to a lower air/fuel ratio.

These BSFC's refer to the engine only and do not include the combustor fuel flow.

3.3.12 Energy Balance Turbocharger -- Take.Off

Turbine pressure ration Prt = 3.052

Compressor pressure ratio Prc = 4.066

Efficiencies: Adiabatic Polytropic Compressor .78 .818 Turbine .77 .744 Mechanical .98 .98 Overall polytropic efficiency S = .596

_F_-= .772

1. Required turbine inlet temperature Tit follows from .286 -4

.875 (Prc Prt = I1 - Tit -1)] Tic Compressor inlet temp. Tic = 273 + 15.5 = 288.5°K Required Tit = 888.9°K

6O 2. Availableturbinetemperature. Thegas in the exhaustmanifold is a mixture of exhaustgasesand scavenge air.

Exhaustgas conditions at exhaustport opening-- Seeparagraph3.3.9.

Ps= 970kPa Ts = 937 + 273 = 1210°K

Thegas expandsto exhaustmanifold pressure(309.5kPa),resulting in a reducedtemperatureof 909.4°K.

Thescavengeair will heat up in the cylinder to 444.4°K. The mixingof exhaustgas and scavengeair results in a mixing temperatureof 806.7°K.

3. The combustormust heat the gas from 806.7°Kto 888.9°Kin orderto provide the turbine energybalance. 4. Theresultantincreasesof the BSFCdue to the combustoroperationare:

A. Takeoff BSFC= 18.2g/kW-hr B. 100%powercruise BSFC = 6.1g/kW-hr

C. 65% powercruise BSFC= 0 g/kW-hr

5. Turbochargerparameters: A. Compressor: Wheeldiameter134.62mm Speeds:

100% Cruise 65% Cruise Take-Off Power Power Shaft RPM 70,052 88,906 83,630 Tip Speed m/sec 494 627 589

B. Turbine: Wheel diameter 127.00 mm

3.3.13 Cooling Requirements

1. Aftercooler:

100% Cruise 65% Cruise Take-Off Power Power Air Flow kg/sec .614 .568 .364 /kt °C 8O.5 116.6 70.0 Heat Rejection kcal/min 721 966 371

61 2. Oil Cooler:

Comparisonwith someexistingenginesof comparablespecificpower: A. VHO-- Caterpillarveryhigh output 8-cylinderwater-cooleddiesel,4-stroke cycle,477kW,oil-cooledpistons.

B. AVCR-1100-- TeledyneContinentalMotorshigh output 12-cylinder, 4-strokecycle,932kW,VCRpistons.

C. GTSIO-520-- 6-cylinderair-cooledgasolineengine,324kW,4-strokecycle.

D. TSIR-5190-- McCulloch,5-cylinderliquid-cooledgasolineengine,2-stroke cycle,201kW.

TABLE XIV Comparative Oil Cooler Data

Heat Rejection Oil Flow Spec. Flow /%-t Engine kW kcallmin kcallmin-kW _ Imin kglmin t Imin-kW kglmin- kW °C VHO 477 806 1.69 208.2 172.1 .44 .36 8.5 AVCR 932 1714 1.84 280.1 231.6 .30 .25 13.4 GTSIO 324 413 1.27 46.6 38.5 .14 .12 19.5 TSIR 201 328 1.63 18.9 15.6 .09 .08 38.1

Weight of oil .827 kg/_..

Spec. heat of oil .55 kcal/kg/°C.

Choice of Aircraft Diesel Oil Cooler Parameters:

A. Spec. heat rejection 1.35 kcal/min/kW which is: • Lower than VHO -- aircraft diesel has no piston cooling jets. • Lower than AVCR -- aircraft diesel has no VCR pistons. • Higher than GTSIO -- aircraft diesel is 2-stroke cycle. • Higher than TSIR -- aircraft diesel has no piston cooling.

Projected heat rejection 298 kW diesel Q = 298 x 1.35 = 403 kcal/min.

B. Spec. flow rate .178 l/min/kW • Lower than VHO -- no oil required for piston cooling. • Lower than VCR -- no oil required for VCR pistons. • Higher than GTSIO -- aircraft diesel is 2-stroke cycle. • Higher thanTSIR-- /kt of the TSIR is too high.

Projected oil flow rate 298 kW diesel 298 x .178 = 53 _/min.

3. Fuel Injectors

Expected heat rejection per injector 25 kcal/min.

62 4. Cooling Requirements:

The total cooling requirements for the 3 modes of operation are shown in Table XV.

TABLE XV Cooling Requirements

100% Power 65% Power Take-Off Cruise Cruise Aftercooler kcal/min 721 966 371 Oil cooler kcal/min 403 403 262 Injectors kcal/min 150 150 98 Total kcal/min 1,274 1,519 731

Fuel flow kg/hr 67.1 65.3 37.8

Heating value 10,250 kcal/kg

Total energy kcal/min 11,463 11,155 6,458

% cooling of total energy 11.1 13.6 11.3

3.3.14 Anticipated Maximum Surface Temperatures of Engine Components

Crankcase 150°C (synthetic oil) Aftercooler (peak) 230°C w/o insulation Compressor housing 230°C w/o insulation Turl_ine housing 595°C w/o insulation (will be radiation shielded) Turbine housing 290°C with insulation Exhaust manifolds 150°C with insulation Combustor surface 150°C with insulation Intake manifolds 95°C Cylinders 150-175°C

3.3.15 Weight of the 298 kW Diesel

A detailed analysis indicates an expected weight of 207.5 kg. This is dry weight and includes all accessories except the filters.

The weight of a comparable gasoline aircraft engine, the GTSIO-520-H is 262.4 kg. Specific component weights are listed in Table XVI.

63 3.3.16Initial Cost of the 298 kW Diesel

The method followed here assumes a certain cost per kg of material. See Table XVI.

TABLE XVI 298 kW Initial Cost -- Aircraft Diesel

A Technology B A x B &/orMat. Weight Eval. Part Reasoning Factor kg No. Prop Gear Housing 1.00 16.06 16.06 Crankshaft 1.00 10.41 10.41 Counterweights Tungsten, high $/kg. 2.00 4.56 9.12 Prop Drive Gears 1.00 9.98 9.98 Crankcase Assy. 1.00 6.54 6.54 Accessory Housing 1.00 3.96 3.96 Accessory Drive Gears 1.00 4.00 4.00 Pistons High technology 5.00 7.48 37.40 Connecting Rods High material cost, partially 1.50 5.62 8.43 offset by simplicity Piston Rings High technology 5.00 .69 3.47 Cylinder Assys. Ceramic liner, otherwise 2.00 34.90 69.79 simplified Injection System Closer tolerances 1.50 5.74 8.61 Intake System 1.00 9.49 9.49 Exhaust Manifold 1.00 12.11 12.11 Fuel Pump 1.00 1.16 1.16 Governor 1.00 .91 .91 Vacuum Pump 1.00 .91 .91 Oil Pumps 1.00 5.95 5.95 Starter/Generator 2.50 6.37 15.93 (total package) Oil Cooler 1.00 2.73 2.73 Aftercooler 1.00 4.55 4.55 Turbocharger High performance 2.00 15.91 31.82 Balance Engine Parts 1.00 37.47 37.47 207.50 310.85

Column A represents a technology and/or material factor which expresses the effect of advanced technology on component cost per kg when compared to current production costs.

Column B shows the calculated weights of components of the proposed engine. The third column then represents the cost ratio of advanced diesel and current technology components.

64 1. Weight Factor:

Weight diesel 207.5kg

Weight gasolineengine 262.4kg Weight ratio 207.5 = .791 262.4

2. Overallenginetechnologyand materialfactor:

TotalAx B _ 310.85 - 1.498 Total B 207.50

3. Overall cost ratio diesel vs. current gasoline engine: .791 x 1.498 = 1.185

3.3.17 Emissions

Emissions were not quantitatively addressed, however, the following qualitative statements are valid:

1. Hydrocarbons and carbon monoxide will be oxidized by the use of a catalytic converter.

2. NOx concentration will be minimized due to the relatively low peak pressures (9,650 kPa) and lower peak temperatures.

3. Smoke levels should be relatively low since the minimum trapped A/F ratio will be on the order Of 24:1.

3.3.18 Noise

As with emissions only, qualitative evaluations were made of the anticipated engine noise as listed below: (propeller noise is covered elsewhere in this report)

1. The catalytic combustor and insulated exhaust stacks in series with the turbocharger should minimize d!rect combustion noise.

2. The absence of cylinder cooling fins should reduce externally generated vibratory noise.

3. The absence of valves, rocker arms, push rods, and camshaft should minimize internally generated mechanical noise.

4. The geared drive will allow a relatively low propeller speed, thereby reducing prop generated noise.

5. Two-stroke cycle operation, however, tends to offset some of the gains noted above.

65 3.3.19 Risk Areas Associated With the Selected Design

Following are the areas where existing technologies need to be advanced to make such an engine feasible':

1. Piston rings -- operating in uncooled cylinders.

2. Cylinders -- ceramic components and their interface with metallic hardware.

3. Turbo starter/alternator operating at high speeds.

4. Catalytic combustor and its associated controls.

5. Cooling of the cylinder exhaust ports.

6. Piston lubrication.

7. Spherical connecting rod end.

Development programs in all these areas are in progress at NASA and TARADCOM (Army).

3.3.20 Proposed Development Program for the 298 kW Diesel Engine

Should the development of such an engine be undertaken, a detailed development program would be recommended based on the following problem areas:

1. Two.Cycle Performance Demonstration.

A. Design and procurement of hardware.

B. Flow modeling.

a. Port configuration

b. Scavenge ratios

c. Timing variations (ports)

d. Air utilization Note: This activity may be deleted. Cost/benefit evaluation in process.

C. Combustion development (SCTE) -- standard cooled cylinder.

a. Piston configurations

b. Injection characteristics • Spray patterns • Timing optimization • Emissions • Smoke • BSFC

66 2. Adiabatic (uncooled) Operation

A. Materials evaluations and selection.

a. Ceramic piston

b. Ceramic cylinder liner

c. Piston ring materials

d. Solid lubricants

B. Design and procurement of hardware.

C. SCTE demonstration.

a. Integrity of ceramic components

b. Demonstration of adequate piston ring sealing and life

c. Port cooling

d. Piston lubrication

e. Injection nozzle cooling

f. Performance • BSFC • Emissions • Smoke • Oil consumption

3. Turbocharger Development

A. Design and procurement of hardware.

B. Compressor bench test.

a. Operation at • High specific speed • High pressure ratios • High flow factors

b. Variable diffuser (if necessary)

c. Maximized efficiencies

67 C. Turbinebenchtest.

a. Operationat • Variableturbinebackpressure(altitude) • Hightip speeds(640m/sec) • TIT815°C • Very high efficiencies

b. Pulse recovery versus steady flow turbine housings

D. Bench test of complete turbocharger.

E. Combustor bench test.

a. Efficiencies

b. Emission control (catalyst)

c. Reliability

d. Life

e. Controls

f. Effect on pulse recovery

4. Support Hardware

A. Design and procurement of hardware.

B. Test of

a. High speed starter/alternator

b. Bleed air starting system

c. Composite rods

d. Injection equipment

e. Synthetic and solid lubricants

f. Catalyst ignition system

g. Electronic controls for • Combustor operation • Injection control • Prop control interface

h. Aftercooler

i. Oil cooler

j. High speed gear train

68 5. Multi-Cylinder Demonstration

A. Design and procurement of hardware

B. Performance and system integration

a. Scavenge characteristics

b. Startability (cold and hot)

c. BSFC

d. Emissions

e. Reliability

f. Altitude operation

C. Demonstrate design integrity

a. Assembly

b. Torsional characteristics

c. Structural integrity

D. FAA type testing

a. Safety

b. Durability

c. Reliability

This type of program could possibly be completed in a 5-6 year time frame and result in a flyable demonstrator engine.

3.3.21 Alternate Technologies

Failure to attain all targets of the development program would not mean a failure of the whole program. The alternate technologies, although less ambitious, will still result in a diesel powerplant which is superior to existing aircraft engines. Some of the alternate solutions are outlined below:

1. Uncooled cylinders:

Alternate: Apply limited cooling to avoid need for ceramics.

The penalties are:

A. Increased fuel consumption (still much lower than gasoline engines).

B. Increased cooling drag.

69 2. High speedalternator: Alternate:Drivethe alternatoroff the engine.

Thepenaltiesare: A. A larger,heavieralternator.

B. Separate,declutchableturbostarterrequired. 3. Catalytic combustor:

Alternate:Conventionalcombustorand ignitor. The penaltiesare:

A. Theignitor must remainturnedon wheneverthe combustoris operating.

B. Emissionslevelof the enginemight be higher.

3.3.22 Comparison of the 298 kW Aircraft Diesel and a Comparable Current Gasoline Engine.

A comparison is made with the 4-stroke cycle GTSIO-520-H gasoline engine.

Table XVII shows this comparison in a tabular form.

Figure 3-30 is a size comparison. The frontal area of the diesel engine is 78% of that of a compatable gasoline engine.

TABLE XVII Comparison of GTSIO-520-H Gasoline and GTDR-290 Aircraft Diesel Engine

4-Stroke Cycle 2-Stroke Cycle GTSIO.520.H GTDR.290 Gasoline Engine Diesel Engine Configuration 6 cyl. opposed 6 cyl. radial Displacement ,_ 8.52 4.71 Take-off RPM 3400 3500 Rated max. take-off power kW 280 298 Rated max. for cruising kW 210 298 Prop speed at take-off RPM 2278 2345 BSFC g/kW-hr: Take-off 425.8 225.1 100% power cruise 219.0 65% power cruise 273.7 194.6 Dimensions: Length mm 1429 1105 Width mm 865 632 Height mm 663 660 Engine weight dry, kg 262.4 207.5

7O .,_ 1429

632

t I 663

I I £-J -I I i I I I / ,,.j I ..,.t I I / L_. r-J r I i I ! I I I i

865 =

GTDR-290 GTSIO-520-H

FIG URE 3-30 SIZE COMPA RISON G TSIO-520-H A ND GTDR-290 AIRCRA FT DIESEL ENGINE 3.4 The 149 kW 4-Cylinder Engine

1. Different design philosophy was applied to the 149 kW engine:

A. To find out how other technologies will affect the configuration and performance of a diesel aircraft engine.

B. To avoid the 149 kW concept from being a scaled down version of the larger powerplant. Chances are that at the conclusion of the development program only one type of configuration will emerge, differing only in size and adaptation details from one engine to another.

C. The technologies applied to the 149 kW engine are not as far advanced as in the case of the 298 kW engine. The 149 kW engine will primarily serve the private owner market where initial cost and ease of maintenance carry more weight than in the case of the corporate aircraft.

D. The engine will be easier to develop and manufacture.

Figure 3-31 shows an a_rtist rendering of the proposed engine.

Figure 3-32 shows the schematic of the engine.

3.4.1 Technologies Applied to the 149 kW Engine

The following features are incorporated in the 149 kW design concept:

1. Radial configuration.

2. Two-stroke cycle Curtis loop scavenging.

3. Minimum cylinder cooling -- reduced fin area.

4. Variable compression ratio pistons (VCR).

5. Mechanically driven centrifugal blower, declutched when not needed.

6. Glow plug starting aid in cylinders.

7. Conventional starter and alternator.

8. Conventional exhaust system (no combustor).

9. Direct propeller drive.

3.4.2 Minimum Cylinder Cooling

Calculations of the heat transfer through cylinder walls and confirmed by tests at TCM/GPD show that the heat flux is highest through the cylinder walls which surround the combustion chamber (when the piston is in top dead center). The maximum gas temperature to which the cylinder wall locally is exposed drops off fast as the piston travels downward, resulting in a lower local average cycle gas temperature and, therefore, a reduced heat flux. It can be safely said that all cooling fins below the piston ring belt (piston in TDC) can be eliminated without an appreciable effect on cylinder wall, piston and piston ring temperatures.

72 FIGURE 3-31 149 kW, 4-CYLINDER DIESEL AIRCRAFT ENGINE INJECTION PUMPS AND NOZZLES

\-- VCR PISTONS f/

REOUCEOCOOLING [ ' TURBINE FUEL PUMP

;NGINEI [ STARTER "_ J____j

ENGINE PROPELLER OIL PUMP GOVERNOR

FIGURE 3-32 SCHEMA TIC 2-STROKE TURBOCHARGED ENGINE REDUCED COOLED/VCR PISTONS

This approach results in an increase of cooling drag when compared to uncooled cylinders but eliminates the need for ceramic components as was recommended for the 298 kW engine.

3.4.3 Variable Compression Ratio Pistons

To keep firing pressures down to 9,650 kPa it is necessary to reduce the compression ratio under load to 10:1. However, the engine cannot be started or run idle at such a low compression ratio. In the case of the 298 kW engine, this was solved by means of the independent turbocharger loop which provides intake air of sufficient pressure and temperature to start the engine and the catalytic combustor which keeps the turbocharger at a high speed during engine idle operation. That is not the case here, therefore for this case a variable compression ratio piston is recommended.

The VCR piston--Figure 3-33, varies the compression ratio from 17:1 at start and low load to 10:1 at full load. This high C.R. is sufficient under normal ambient conditions to start the engine.

74 I OIL -- X

FIGURE 3-33 VARIABLE COMPRESSION RA 7"10PISTON

&4.4 Mechanically Driven Centrifugal Blower

Scavenging of a 2-stroke cycle cylinder requires that the intake manifold pressure exceeds the exhaust manifold pressure at any load and engine speed. The turbocharger, however, produces a negative /kp at low load. This is no problem for 4-stroke cycle engines where the piston does the scavenging. The 2-stroke cycle engine without a combustor requires an engine driven blower to produce a positive Z_p across the cylinders at low loads. The blower will be disconnected at the load point where the turbocharger provides a positive Ap.

3.4.5 Glow Plug Starting Aid in Cylinders

Even the 17:1 compression ratio does not provide a sufficiently high compression temperature to ignite the fuel at very low ambient temperatures. Operation of the glow plug may be required tb assure good startability. It is also intended that glow plug operation would automatically be in effect at low throttle settings. This would be an added safety feature to assure absolutely no misfiring during descent mode operation.

3.4.6 Direct Propeller Drive

1. It became obvious early in the design phase of the 149 kW engine that a direct drive would result in a smaller engine package and a weight reduction.

2. The engine reliability is improved by this approach due to fewer parts.

75 3.4.7 Initial Performance Parameters

The chosem BMEP of approximately 1200 kPa is 100 kPa higher than the BMEP of the 298 kW engine. The much lower crankshaft speed dictated by the direct propeller drive will result in better scavenging and, hence, a larger amount of air trapped in the cylinders. We should, therefore, be able to obtain a higher BMEP without an increase of cylinder temperatures. The detailed cycle calculations (temp. T4 of the p-V diagrams) bear this out. A stroke/bore ratio = 1 was chosen.

The engine characteristics become:

Number of cylinders 4 Take-off power 149 kW Engine speed at T.O. 2400 RPM BMEP 1187 kPa Displacement 3.14 liter Cylinder bore 100 mm Stroke 100 mm Piston speed 8.00 m/sec Propeller drive Direct

The piston speed is very comfortable and will result in long piston ring life.

3.4.8 Engine Concept Design

The engine concept design is shown in the Figures 3-34 through 3-38. The cylinders are arranged in one bank of four cylinders. The rotating and reciprocating inertias are 100% balanced. The cylinders have a limited number of cooling fins to cool the combustion chamber. The necessity for a gear driven blower at the back side of the engine made it more practical to have the cylinder intake port at the back side and the exhaust manifolds at the front. The exhaust manifolds will be insulated to avoid radiation to the injection pumps. Two exhaust manifolds are required to avoid pulse interference between cylinders. The connecting rods are executed as slipper rods. The big ends of the rods are wider than in the case of the 298 kW engine to compensate for the reduced circumferential contact length.

The use of synthetic oil is not contemplated because of lower cylinder temperatures but may be feasible to extend the periods between oil changes. Four individual injection pumps are provided driven off a single lobe cam ring. The centrifugal blower is driven off the propeller shaft through a lay shaft which is located above the crankcase between the cylinders #1 and #4. This arrangement was chosen rather than a drive from the rear end to avoid torsional problems. The nodal point lies close to the largest inertia member of the crankshaft system, that is the propeller. Putting the blower drive gear near this point reduces the input of torsional amplitudes into the blower drive. The lay shaft, which is a quill shaft, further isolates the blower from the crankshaft vibrations. However, this feature also necessitates a direct propeller drive. To put a propeller reduction gearing in front of the blower drive would have led to an unacceptable length of the engine. A weight analysis for this particular engine showed that the direct drive with the inherent larger piston displacement results in a lighter engine than the indirect drive.

76 EXHAUST MANIFOLDS _XX. CENTRIFUGAL •COMPRESSOR MAGNETIC CLUTCH \ DISC CLUTCH i i

I L

\

INJECTOR PUMPS EXHAUST PORT COOLING 7 , PENDULUM

_L_ DAMPER

EXHAUST" L___.'_>,_j/_ INTAKE PORT VCR PORT _'_j__'- -"_-_ PISTONS SECTION A-A (VIEW THRU EXHAUST & INTAKE PORTS)

FIGURE 3-34 149 kW AIRCRA FT DIESEL--LONGITUDINAL SECTION GLOW PLUGS

INJECTORS

OIL SCREEN VACUUM PUMP OIL PUMP

STARTER

QO

FIGURE 3-35 149 kW AIRCRA FT DIESEL--CROSS SECTION INJECTORS 1

OIL SCREEN

ALTERNATOR OIL PUMP VACUUM PUMP-_ 668.02 (26.32)

STARTER_

341.376 (13.44)

335.28 (13.20)

657.28 (25.88)

FIGURE 3-36 149 kW AIRCRAFT DIESEL--FRONT VIEW FLEXIBLE CONNECTION

EXHAUST MANIFOLDS/ INTAKE MANIFOLD TO CYLINDER // 938.8 (37.00) / 4434 (1746) , '! INJECTOR PUMPS MAGNETIC CLUTCH (BLOWER DRIVE) TURBOCHARGER

FUEL TRANSFERPUMP

CO O

GOVERNOR VARIABLE TURBINE NOZZLE ACTUATOR

OIL TO SCAVENGE PUMP

FIGURE 3-37149 kW AIRCRAFT DIESEL--SIDE VIEW ENGINE MOUNTING INTAKE MANIFOLD BOSSES

TURBO EXHAUST GASES _

TACHOMETER PAD TURBO AIR INLET AND 20005

¢O

OIL COOLER AFTERCOOLER 190 (7.50)

FIGURE 3-38 149 kW AIRCRAFT DIESEL--REAR VIEW CENTRIFUGAL EXHAUST MANIFOLDS COMPRESSOR

TURBOCHARGER MAGNETIC CLUTCH DISC CLUTCH

I \ VARIABLE TURBINE NOZZLE ACTUATOR INJECTOR PUMPS EXHAUST PORT DAMPER COOLING / AFTERCOOLER I/__NTAK E PORT EXHAUST VCR OIL TO SCAVENGE PORT PISTONS PUMP SECTION A-A [VIEW THRU EXHAUST & INTAKE PARTS]

FIGURE 3-39 149 kW AIRCRAFT DIESEL--LONGITUDINAL SECTION GLOW PLUGS

INJECTORS

OIL SCREEN VACUUM PUMP \ OIL PUMP

¢O STARTER U_

i! OIL COOLER I

"_AFTERCOOLER

FIGURE 3-40 149 kW AIRCRAFT DIESEL--CROSS SECTION INJECTORS

OIL SCREEN

ALTERNATOR OIL PUMP VACUUM PI. 608.55 [23.88]

STARTER oo \_

317.5 [12.50]

AFTERCOOLER "_

- OIL COOLER

322 335.28 _! [12.68] 8001 -- [13.20] [31.50]

FIGURE 3-41 149kW AIRCRAFT DIESEL--FRONT VIEW FLEXIBLE CONNECTION INTAKE MANIFOLD TO CYLINDER 965.2 4434 [3.8.00] [17.46] EXHAUST MANIFOLDS MAGNETIC INJECTOR

CLUTCHDRIVE] PUMPS

[BLOWER FUEL \ / TRANSFER PU

r 00 U1

C.G. [.02] =BLE TURBINE NOZZLE 1 ACTUATOR I GOVERNOR

OIL TO SCAVENGE PUMP /

FIGURE 3-42 149 kW AIRCRAFT DIESEL--SIDE VIEW ENGINE MOUNTING INTAKE MANIFOLD BOSSES

TURBO EXHAUST GASES

TACHOMETER PAD TURBO AIR INLET AND 20005

00 O_

ij STARTER

289.05 [11.38] ] / t AFTERCOOLER OIL COOLER 464.82 [18.3]

FIG URE 3-43 149 k W A IRCRA F T DIESEL-- REA R VIE W The blower drive is provided with two clutches. One, the magnetic clJtch, disengages the blower drive once the turbocharger has come up to speed. The location of the magnetic clutch is such that as much of blower drive as possible is disengaged to prevent unnecessary drag on the engine. A disc type slip clutch is provided to prevent large torsional amplitudes as they occur at low engine speeds due to cyclic irregularity from reaching the blower. The turbocharger is mounted behind the engine as are the oil cooler and the aftercooler.

A second version of the engine was drawn, Figures 3-39 through 3-43, which accommodates a retractable nose gear. The coolers are moved outboard and the turbocharger is raised to provide space between Cylinders #2 and #3 for the nose gear strut. The increased width of the engine is no problem since it occurs near the fire wall where the width of the fuselage is determined by the side-by-side cabin seating arrangement.

3.4.9 149 kW Engine Operating Data

The following operating parameters have been calculated for the 149 kW engine:

TABLE XVlll Engine Operating Parameters

100% Power 65% Power Take-off Cruise Cruise Altitude 0 3,048 3,048 meters Power 149 149 97 kW R P M 2400 2400 1800 Displacement 3.14 3.14 3.14 liters Bore x Stroke 100 x 100 100 x 100 100 x 100 mm BMEP 1,187 1,187 1,029 kPa Compressor Pressure Ratio 4.16:1 6.10:1 4.13:1 Compression Ratio Variable Variable Variable Max. C.R. 17:1 (effective) Min. C.R. 10:1 (effective) Barometric Pressure 101.4 69.6 69.6 kPa Ambient Temperature 15.5 - 5 - 5 °C Intake Manifold Pressure 411.8 411.8 280.9 kPa Intake Manifold temperature 116 116 116 °C Exhaust Manifold Pressure 316.8 316.8 255.4 kPa Scavenge System Curtis Loop Curtis Loop Curtis Loop Scavenge Ratio 1.3 1.3 1.3 Ratio Boost/Backpressure 1.3 1.3 1.1 Height Intake Ports 20.13 20.13 20.13 mm Height Exhaust Ports 27.15 27.15 27.15 mm Intake Ports Open/Close - 61 ° _ 61 o _+61 ° BBDC/ABDC Exhaust Ports Open/Close _+71 ° _+71 o + 71 ° BBDC/ABDC BSFC 222.0 228.1 209.8 g/kW-hr. Fuel Flow 33.1 34.0 203 kg/hr. Air/Fuel Ratio 26.6 26.0 24.0

87 3.4.10 P-V Diagrams

Following are calculated air cycle performance data for the proposed 149 kW engine. Figure 3-21 illustrates the points calculated on the P-V diagram.

Compression pV 1.37 = C

Expansion pV 1.285 = C TABLE XIX Air Cycle Performance

100% Power 65% Power Take-off Cruise Cruise

P_ 364 364 268 kPa Vl .637 .637 .623 liter Ti 149 149 138 °C P2 8,540 8,540 8,540 kPa V2 .064 .064 .050 liter T2 717 717 773 °C P3 9,650 9,650 9.650 kPa V3 .064 .064 .050 liter T3 846 846 911 °C P, 9,650 9,650 9,650 kPa V4 .118 .120 .095 liter T4 1,803 1,829 1,975 °C Ps 1,110 1,120 830 kPa V5 .637 .637 .623 liter T5 1,012 1,032 1,032 °C Fuel/Cyl./Rev. .0000576 .0000590 .0000472 kg Air in Cylinder .00192 .00192 .00142 kg Q/Cyl./Rev. .590 .605 .484 kcal Q, .052 .052 .041 kcal Q2 .538 .552 .443 kcal IM EP 1,539 1,582 1,416 kPa Mech. Eff. (engine) 77 75 73 % Turbine Press. Ratio 3.123 4.549 3.667 Compressor Pressure Ratio 4.156 6.101 4.126 Compressor Efficiency 81.5 80.5 79 % Turbine Efficiency 80.5 79.5 78 % Mechanical Efficiency 98 98 98 % Overall Turbo Efficiency 64.3 62.7 60.4 % Required TIT 352 362 374 °C Fuel Flow 33.1 34.0 20.3 kg/hr. Air Density .00301 .00301 .00227 kg/,_ FLOWS: Weight Pure Air .245 .245 .142 kg/sec Weight Fuel .009 .0095 .006 kg/sec Weight Exhaust Gas .254 .2545 .148 kg/sec Weight Scavenge Air .073 .073 .043 kg/sec

The Figures 3-44 through 3-46 show the schematics of the three operating conditions.

The engine performance curves are shown in Figure 3-47.

88 .318 kg/sec

BLOWER INTERCOOLER • COMPRESSOR AIR CLEANER

411.8 kPa

115.6°C ID--I ! I I I I 411.8 kPa I 416.4 kPa 100.1 kPa I I I 115.6°C 193.9°C Prc= 4.156 } 15.5°C I 101.415.5°CkPa I I I L_J

¢o

316.8 kPa 101.4 kPa

552°C I Prt=3.123 J

TURBINE

149 kw

w 2400 RPM .327 kg / sec SEA LEVEL

FIG URE 3-44 OPERA TING SCHEMA TIC-- TA KE-OFF .318 kg/sec

BLOWER INTERCOOLER COMPRESSOR AIR CLEANER

411.8 kPa ,-3 115.6°C I i i I 411.8 kPa 420.7 kPa Prc = 6.101 I I .50C -5°C i I 115.6*C 220.6°C I I I L__I

O

; 316.8 kPa

561.7°C Prt=4.549 69.6 kPa

TURBINE

v 149 kW .328 kg/sec 2400 RPM 3,048m

FIG URE 45 OPERA TING SCHEMA TIC-- 100% CRUISE PO WER .185 kg/sec

BLOWER INTERCOOLER COMPRESSOR AIR CLEANER

280.9 kPa 115.6°C i-7 i I 69.0 kPa 69.6 kPa I I 280.9 kPa 284.5 kPa I t t Prc=4.126 ,J .5oc .5oc I I 115.6°C 165.0°C I I I__1

_D

255.4 kPa 69.6 kPa 573.9°C Prt=3.667 I

TURBINE

v .191 kg/sec 97 kw 1800 RPM 3,048m

FIGURE 3-46 OPERA TING SCHEMA TIC--65% CRUISE PO WER ENGINETORQUEAT 140 640 CRANKSHAFTSPEED E 620 la.J 130 6O0

I.,-, 580 120

110 LOADPOWER

100

90

he, LL,

a,. 80 LOADPOWER

70

60

-J .,-I 50 BSFC

40

I I I I 1400 1600 1800 2000 2200 2400 ENGINERPM

FIGURE 3-47 SEA LEVEL PERFORMANCE 4-CYLINDER RADIAL AIRCRAFT DIESEL ENGINE

92 3.4.11 Stress Calculations

All stress calculations are based on a 9650 kPa firing pressure. Figure 3-48 shows the cylinder arrangement. The cylinders are in one plane. The firing order is 1,2, 3, 4 with even 90 ° firing intervals.

1 F]

F] 4

FIGURE 3-48 CYLINDER ARRANGEMENT

1. Power Train Data

Weight piston assembly: Ring carrier assembly .93 kg Pin carrier assembly 1.86 kg Total piston 2.79 kg

Composite connecting rod .33 kg Slipper rings .19 kg Counterweights 7.70 kg

Total reciprocating WR 27.9 kg-cm Total rotating WR 20.8 kg-cm Total counterweight WR 47.1 kg-cm

Balance 100%

93 2. Crankshaft Stresses

A. Crankpin Fillet Radius Max. principal bending stresses 584 MPa Min. principal shear stresses 128 MPa

B. Power Transmission Shaft Nominal shear stress 174 MPa

C. Material AMS 6415 Ultimate tensile strength (min) 1034 MPa Endurance strength (machined & peened) 552 MPa

D. B!ower Quill Shaft Max. power transmission capacity 39 kW

3. Connecting Rod Stresses and Bearing Pressures

A. Connecting Rod Max. compressive stress 214 MPa Min. compressive stress 27 MPa Composite material fatigue strength 391 MPa

B. Crankpin Bearing Unit Load 28 M Pa SAE 794 leaded bronze_max, unit load 69 M Pa

C. Piston ball joint (30 mm_) unit load 91 MPa (Note: w/o oil groove on ball)

g. Main Bearing (55 mm_ x 28 mm length) Peak unit load 21 MPa Min. unit load 6 MPa

The Figures 3-49 and 3-50 show the main bearing load diagram and the crankshaft and connecting rod stresses.

4. Cylinder Barrel Stresses

A. 8-Cylinder Hold Down Studs M10X1.5 -- 6g Grade 8 (proof load 40,430 N/stud) Torque to 75% proof load 30,320 N/stud Peak dynamic load: 1,490 N/stud

B. CylinderWall Hoop Stresses 63 MPa

C. CylinderWall Longitudinal Stresses 32 M Pa

D. Material: Steel Min. flexural strength at 1000°C 455 MPa

94 210" 150" \

140 _ 140" 220"

230" 130' 130 ° 230'

240" 120' 120 ° 240'

] 250 ° 110' 110 = 250'

260" 100' 100 ° 260'

270" 90' 90" 270'

, 280 ° 80" 80 ° 280'

70' 290 °

, 300 ° 60 ° J 300"

310 ° 50" 50 ° , 310"

-q

40 ° i 320.40 ° 320" J 5

_ J 330" ' 3o" 3_° 350° 0: mo 20" 20" 10 ° 350" 340"

FIGURE 3-49 MAIN BEARING LOAD 149 kW AIRCRA FT DIESEL

95 AMS.6415 1034 MPaUTS MPa 1000-

C'RKPIN FILLET [2300 RPM]

'4 -200 400 600 800 1000 MPa MEANSTRESS

CON-ROD[2300 RPM]

-600

FIGURE 3-50 CRANKSHAFT AND CONNECTING ROD STRESS

96 5. Natural Frequencies Crankshaft System

A. First Mode's Natural Frequency 16 Hz

1st Order 974 rpm

B. 2nd Mode's Natural Frequency 78 Hz

1st Order 4655 rpm 2nd Order 2327 rpm 4th Order 1164 rpm

Pendulum dampers to be tuned for 2nd and 4th orders.

3.4.12 Projection of Fuel Consumption

Reference: Paragraph 3.3.11.

Baseline for th'e projections of BSFC was the 2-stroke cycle 8V92T Detroit Diesel Allison engine:

Min. BSFC = 229 g/kW-hr

In the case of the 298 kW engine a 15% BSFC improvement was projected due to the uncooled cylinders, resulting in a 65% cruise power BSFC (best economy) of .85 x 229 = 194.6 g/kW-hr.

For the 149 kW engine with partially cooled cylinders a 8.5% gain is projected resulting in a minimum BSFC = .915 x 229 = 209.8 g/kW-hr BSFC.

Projected BSFC's g/kW-hr:

298 kW Engine 149 kW Engine 65% Cruise Power 194.6 209.8 Take-off Power 194.6 + 12.2 = 206.8 209.8 + 12.2 = 222.0 100% Cruise Power 194.6 + 18.3 = 212.9 209.8 + 18.3 = 228.1

These values were used in Paragraph 3.6.9.

3.4.13 Cooling Requirements

Reference: Paragraph 3.3.13

1. Aftercooler

100% Cruise 65% Cruise Take.off Power Power Air Flow .3t8 .318 .185 kg/sec _t 78.3 105.0 49.4 °C Cp .243 .243 .243 kcal/kg- °C Heat 363 487 133 kcal/min

97 2. Oil Cooler

Specific heat rejection of the VHO is 1.69 kcal/min/kW

Heat at 149 kW: Q = 149 x 1.69 = 252 kcal/min.

100% Cruise 65% Cruise Take-off Power Power Q kcal/min 252 252 164

3. Cylinders

Reference: Paragraph 3.3.11

A. Fully Cooled AVCR-1360: Q = 6,300 kcal/min at 1120 kW Spec. heat 5.625 kcal/min/kW

B. Limited Cooled TDR-192 aircraft diesel: 40% reduction of cylinder heat load Q = .60 x 5.625 x 149 = 504 kcal/min

100% Cruise 65% Cruise Take-off Power Power Q kcal/min 504 504 327.6

The heat balance is shown in Table XX. Figure 3-51 shows a heat balance comparison of the 6-cylinder uncooled, and the 4-cylinder partially cooled aircraft diesels.

A comparison of the Figures 3-51 b and 3-51-c shows that in the case of the 298 kW a large portion of the cooling loss reduction ends up in the exhaust gases. This is why Cummins Engine Company decided on turbocompounding of their adiabatic diesel engine to utilize this energy. Turbocompounding of the aircraft diesel was rejected for the following reasons:

a) Increased weight of the powerplant.

b) Reduced reliability of the high speed gear train.

The penalty is a somewhat higher BSFC.

3.4.14 Anticipated Maximum Surface Temperatures of Engine Components

Crankcase 150°C Aftercooler (peak) 195°C w/o insulation Compressor housing 195°C w/o insulation Turbine housing 510°C w/o insulation (will be radiation shielded) Turbine housing 230°C with insulation Exhaust manifolds 150°C with insulation Intake manifolds 95°C Cylinder 230-245 ° C

98 TABLE XX

Heat Balance 149 kW Engine

100% POWER 65% POWER TAKE-OFF CRUISE CRUISE

Aftercooler kcal/min 363 487 133

Oil cooler kcal/min 252 252 164

Cylinders kcal/min 504 504 327.6

Total cooling kcal/min 1,119 1,243 625

Fuel flow kg/hr 33.1 34.0 20.3

Total energy kcal/min 5,667 5,812 3,475

Engine power kcal/min 2,134 2,134 1,389

Exhaust gas kcal/min 2,244 2,261 1,357

A. 100% CYLI DER COOLING

B. 60% CYLINDER COOLING C. NO CYLINDER COOLING 149 kW DIESEL 298 kW DIESEL 100% CRUISE POWER 100% CRUISE POWER

FIGURE 3-51 HEA T BALANCE COMPARISON

99 3.4.15 Turbocharger Operation

The turbocharger data for the 3 modes of operation are shown in Table XXI.

3.4.16 Blower Operation

Assume crank-up to 600 rpm. Required pressure ratio across cylinder ports is 1.01. Required blower tipspeed 41.1 m/sec. Blower diameter 177.8 ram.

Blower speed is 4500 rpm at 600 rpm crankshaft speed.

Temperature increase in blower:

AT = .289 (1,01 .286 -1) = 1.4°C .60

Intake manifold temperature 16.9°C

Temperature at the end of compression in the cylinder:

T = (273 + 16.9) x 17.37 = 8270R = 554oc

Fuel ignition temperature is approximately 590°C, therefore, glow plugs are required for start and restart.

3.4.17 Weight of the 149 kW Diesel

A detailed analysis indicates an expected weight of 163.2 kg. This dry weight includes all accessories.

Specific component weights are listed in Table XXlI.

The weight of a comparable gasoline aircraft engine, the TSIO-360-E is 174.6 kg.

3.4.18 Initial Cost of the 149 kW Diesel

Method followed is the same as described in Paragraph 3.3.16. See Figure 3-52.

1. Weight Factor: Weight diesel 163.2 kg Weight gasoline engine 174.6 kg Weight ratio 163.2 _ .935 174.6

2. Technology and Material Factor: Total A x B = 212.71 = 1.303 Total B 163,20

_. 3. Overall Cost Ratio Diesel vs. Current Gasoline Engine:

.035 x 1.303 = 1.218

_ 100 TABLE XXI Turbocharger Data 100% Power 65% Power Take-off Cruise Cruise

Compressor Ratio Prc 4.156 6.101 4.126 Turbine Ratio Prt 3.123 4.549 3.667

EFFICIENCIES: -- Compressor .842 .834 .817 -- Turbine .775 .768 .752 -- Mechanical .980 .980 .980 -- Overall .640 .627 .604 Air Flow .318 .318 .185 kg/sec Exhaust Gas .327 .328 .191 kg/sec Compressor Inlet Press. 101.4 69.6 69.6 kPa Compressor Inlet Temp. 15.5 - 5 - 5 °C °C Compressor Discharge Temp. 193.9 220.6 165.0 Turbine Inlet Temp. 552.0 561.7 573.9 °C Mechanical Blower Operation no no no Exducer Diam. 76.2 76.2 76.2 mm Turbine Rotor 101.6 101.6 101.6 mm Compressor Wheel 114.3 114.3 114.3 mm N/,J-_- 79,170 89,840 77,515

3.4.19 Emissions

. Hydrocarbons and carbon monoxide will be comparable to current 2-stroke cycle engines. A catalytic converter may be added downstream of the turbocharger if future regulations mandate lower HC and CO levels.

2. NOx concentration will be minimized due to the relatively low peak pressures (9,650 kPa) and lower peak temperatures.

3. Smoke levels should be relatively low since the minimum trapped A/F ratio will be on the order of 25:1.

3.4.20 Risk Areas Associated with the Selected Design

Following are the areas where existing technologies need to be advanced:

1. VCR piston -- develop for 2-stroke cycle operation.

2. Cylinders -- reduced cooling air flow.

3. Piston rings -- operating in reduced cooled cylinders.

4. Cooling of cylinder exhaust ports.

5. Piston lubrication.

6. Spherical connecting rod end.

101 TABLE XXII Initial Cost 149 kW Aircraft Diesel

A Technology B A x B &/or Mat. Weight Eval. Part Reasoning Factor kg No. Crankshaft 1.00 4.50 4.50 Counterweights Tungsten 2.00 6.68 13.36 Quill Shaft 1.00 2.71 2.71 Prop Shaft 1.00 4.95 4.95 Pistons 2-Stroke cycle VCR 2.50 10.75 26.86 Piston Rings Elevated temperature 2.50 .24 .59 Connecting Rods Composite material 1.50 1.92 2.88 Cylinders Limited cooling 1.50 29.67 44.51 Injection System Tight tolerances 1.50 3.58 5.38 Front Acc. Gears 1.00 3.08 3.08 Front Housings 1.00 10.57 10.57 Blower Drive 1.20 5.25 6.30 Crankcase 1.00 5.35 5.35 Rear Acc. Gears 1.00 1.57 1.57 Intake System 1.00 5.58 5.58 Blower 1.00 .66 .66 Exhaust Manifolds 1.00 .66 .66 Aftercooler 1.00 3.76 3.76 Turbocharger 1.50 15.45 23.17 Oil Pump 1.00 5.46 5.46 Vacuum Pump 1.00 .91 .91 Governor 1.00 .91 .91 Alternator 1.00 4.88 4.88 Fuel Pump 1.00 1.16 1.16 Starter 1.00 3.54 3.54 Oil Cooler 1.00 2.35 2.35 Balance Parts 1.00 27.06 27.06 163.20 212.71

3.4.21 Proposed Development Program for the 149 kW Diesel Engine

Following is a detailed program that is recommended for the development of this engine:

1. 2-Cycle Performance Demonstration

A. Design and procurement of hardware.

B. Flow modeling.

a. Port configuration

b. Scavenge ratios

c. Timing variations (ports)

102 d. Air utilization Note:Thisactivity maybe deleted.Cost/benefitevaluationin process. C. Combustiondevelopment(SCTE)-- standardcooled cylinder. a. Piston configurations b. Injection characteristics • Spraypatterns • Timingoptimization • Emissions • Smoke • BSFC

2. ReducedCylinderCoolingOperation A. Materialsevaluationand selection.

a. Codling fins configuration b. Piston ring materials c. Solid lubricants

B. Designand procurementof hardware. C. SCTEdemonstration.

a. Cylinderintegrity

b. Demonstrationof adequatepiston ring sealingand life c. Port cooling d. Piston lubrication

e. Injection nozzlecooling f. Performance • BSFC • Emissions • Smoke • Oil consumption

3. VCR Piston Development

A. Design and procurement of hardware.

B. Bench test.

103 C. SCTEdemonstration.

a. Performancetests

b. Endurancetesting

4. TurbochargerDevelopment A. Designand procurementof hardware.

B. Compressorbenchtest. a. Operationat: • High specific speed • High pressureratios • High flow factors b. Maximizedefficiencies C. Turbinebenchtest.

a. Operationat: • Variableturbine back pressure(altitude) • Hightip speeds • Highefficiencies b. Optimizationof variablenozzleareaoperation c. Developmentof nozzlecontrol actuator

D. Benchtest of completeturbocharger 5. SupportHardware A. Designand procurementof hardware. B. Test of:

a. Compositerods

b. Blowerdriveand declutch system

c. Injectionequipment d. Lubricants e. Electroniccontrols for • VATturbocharger • Blowerdeclutching • Injectioncontrol • Propcontrol interface

104 f. Aftercooler

g. Oil cooler

6. Multi-Cylinder Demonstration

A. Design and procurement of hardware.

B. Performance and system integration.

a. Scavenge characteristics

b. Startability (cold and hot)

c. BSFC

d. Emissions

e. Reliability

f. Altitude operation

C. Demonstrate design integrity.

a. Assembly

b. Torsional characteristics

c. Structural integrity

D. FAA Type Testing.

a. Safety

b. Durability

c. Reliability

3.4.22 Comparison of the 149 kW Aircraft Diesel and a Comparable Current Gasoline Engine

A comparison is made with the 4-stroke cycle TSIO-360-E gasoline engine.

Table XXIII shows this comparison in a tabular form.

Figure 3-52 is a size comparison.

105 TABLE XXIII Comparison of TSIO-360-E Gasoline and TDR-192 Aircraft Diesel Engine

4.Stroke Cycle 2-Stroke Cycle TSlO-360.E TDR-192 Gasoline Engine Diesel Engine Configuration 6 cyl. opposed 4 cyl. radial Displacement _. 5.91 3.14 Take-off RPM 2800 2400 Rated max. take-off power kW 149 149 Rated max. for cruising kW 112 149 Prop drive direct direct BSFC g/kW-hr: Take-off 377.1 222.0 100% power cruise _ 228.1 65% power cruise 267.6 209.8 Dimensions: Length mm 1188 965 Width mm 795 800 Height mm 672 607 Engine weight dry, kg 174.6 163.2

4 1188

\\ 795 965 /

! / J

/J

l i -%% - 607 \ \ \

672

_f_ I I.... L__ I

8OO TDR-192 .... TSIO=360-E

FIGURE 3-52 SIZE COMPARISON TSIO-360E AND TDR-192 AIRCRAFT DIESEL ENGINE

106 4.0 ENGINE/AIRFRAME INTEGRATION

This study was conducted by Beech Aircraft Corp. to evaluate the integration of the proposed diesel aircraft engines into future airframes and to determine the effect of the engine on aircraft performance and operating costs. The results were then compared with corresponding data for current production type gasoline engine powered aircraft.

4.1 Engine Installation

Installation design layouts were made which show the 298 kW diesel--Figures 3-16 through 3-20--mounted on a twin engine airplane and the 149 kW engine, Figures 3-39 through 3-43, installed in a single engine aircraft with retractable landing gear.

The Figures 4-1 through 4-3 show the twin engine installation, the Figures 4-4 through 4-6 show the single engine installation.

4.1.1. Description of the layouts

. Engine mounts are of two basic types -- cantilever and bed mount. A cantilever mount from the firewall was used in the twin and a bed mount incorporating the nose gear support structure was used in the single. "Dynafocal" type mounts would be used with the cantilever method to minimize vibration transmission to the airframe.

. The induction system in both cases would be a NACA flush inlet, ducting and an air filter. Alternate air would be available to the engine through a door operated by differential pressure.

. Both engines have a dry oil sump and require external oil tanks mounted in the engine compartments.

OIL SYSTEM DATA

Oil Flow Oil Capacity -- Liters Engine llmin. Engine Sump Tank Total 298 kW Diesel 53 4 15.5 19.5 149 kW Diesel 34 2 9.5 11.5

o Both engines would have cooling air inlets providing air to a plenum chamber. Ducts from the plenum would direct air to individual cylinders, oil coolers, aftercoolers and fuel injectors as needed. On the single, cooling air exits are outboard of the nose gear on the lower side of the cowling. Exits from the twin nacelle would be at the lower aft end.

. The installation drawings were done in enough detail to indicate the features noted above and to provide reasonable assurance that no major installation problems would be encountered with the proposed diesel engine concepts.

107 I SECTION A-A A

J O COMBUSTION AIR PLENUM CO

FUEL PUMP

COOLING AIR IN

OIL TANK

\ \ \ INJECTOR COOLING AIR \ A

COOLING AIR IN AFTERCOOLER COOLING AIR EXIT DUCTED THRU TWO INLETS AFTERCOOLER "_'- COOLING AIR EXIT

FIGURE 4-1 SIDE VIEW TWIN INSTALLATION COOLING AIR INLET INJECTORS [TYP. 6 PLACES]

AIR INLET LER

_.AFTERCOOLER OIL ( AIR INLET AFTERCOOLER i

FIGURE 4-2 FRONT VIEW TWIN INSTALLA TION

109 COOLING AIR IN OIL COOLER

COOLING AIR INJECTOR \ COOLING AIR EXIT OIL COOLER COMBUSTION _ INLET

,. -, --/-- _+-_--_ ' " / COOLER _ I L.AIR FILTER /" .I-- / ' I _ I "___ "',

i i : i : I __7"_.-----_-: I t _ //" !ti COMBUSTION ',, /!_L h ....-i----_--q-_l I __-_-: :_/ I_l 'AIR PLENUM I _"t_ : ' --"_' ..... ' ' i COOLINGAIR i _--- -- t--£ _ __1---1- ' I itil IPU 'PLENUM Iil --- , .__ ___l .... i..L__ _ r..._ _ ltlj-,- ...... I ' J "'-- m r.... , ----- I _"-7 I I ; '_----'- ./' OIL TANK r _ _ i I t _ ------_-- -- - USABLE 4 GAL. z -i- - i /--17 AFTERCOOLER' _ J _ EXP. SPACE 3 GAL.

_-_ COMBUSTION EXHAUST / COOUNGA..,NAFT.COO'. - / -- 2 2

COOLING AIR EXIT AFTERCOOLER

FIGURE 4-3 TOP VIEW TWIN INSTALLATION .ii

FIGURE 4-4 SIDE VIEW SINGLE INSTALLATION m ......

:)LING AIR INLET

COOLING AIR INLET

AFTERCOOLER COOLING AIR INLET STARTI OIL COOLER f \ L\ I ', L\ ' " /

[...... _-.-_-. _ l i . I _ / AFTERCOOLER- - _ I _ i _ _ / \ '% -;...._ -- _-_OILCOOLER/ /

T COMBUSTION EXHAUST

FIGURE 4-5 FRONT VIEW SINGLE INSTALLA TION

112 -'_l COMBUSTION AIR INLE___

/coo.,.o.,__ __!__--_i l]Ji...coo.. |AFTSRCOOLER--_'__T ER'_L"_--"_ ._ _ !_' ' _ i HEAOANOCYL,.OE.____ICOOLINGAIRF_---/ ,i ilk._,_l ---"__" I!'J_1:1' - ' ' _:/I I' ' ,/l: _--J __ ....-_--_---_-_;---_-'_"

i - _-r ...... __ ...... =___ ...... -

_'j_ COOLING-=[ _ ; -; , --,_ --It ......

COOLING AIR '_ "i-_ "_--_---- -!- _! ..... i r J _ HEAOANOCYLINDER L"-_tll _ '_ OIL COOLER "_// _'--_ / O,LCOOLER-- _ L J ' f ...... I / "_ ',..... ----T-,, ----, "-_ I '

i

FIGURE 4-6 TOP VIEW SINGLE INSTALLATION 4.2 Aircraft Configurations

Three view sketches of the airplanes are shown in the Figures 4-7 and 4-8. Following are some characteristics of both planes:

4.2.1 Twin Engine Airplane

Figure 4-7 shows the twin engine aircraft. The sketch yields the following information:

1. Propeller Data

Prop. diameter 2.057 m Prop. speed at take-off 2,345 rpm Tip speed at take-off 253 m/sec = .74a a = Velocity of sound = 20.06 x/_ m/sec (T in °K) At standard ambient temp. 15.5°C a = 20.06 _/273 + 15.5 = 341 m/sec Prop. speed at economy cruise 1,790 rpm Tip speed at economy cruise 193 m/sec Prop. ground clearance 330 mm

2. Sight Angles

The pilot's sight angles for the twin are indicated by A and B (Figure 4-7). The centerline angle over the nose, A, as indicated is about 12 ° . If the airplane were lofted, the angle from the pilot's actual eye position would be about 18° which is considered more than adequate. The smallest lateral angle 8 is 10 ° . This is also more than adequate, especially compared to some current piston engine twins with larger nacelles.

3. Aircraft Data

Twin Engine Twin Engine Diesel Gasoline Airframe minus engine (a) kg 1,860 1,860 Engines (2) Figure 3-36 (b) kg 415 525 Empty weight (a) + (b) (c) kg 2,275 2,385 Payload (d) kg 726 671 Fuel load (e) kg 653 598 Useful load (d) + (e) (f) kg 1,379 1,269 Max. take-off weight (c) + (f) kg 3,654 3,654 Wing span m 13.05 13.05 Length m 11.89 11.89 Tail height m 3.87 3.87 Tail span m 5.09 5.09 Wing area m 2 22.39 22.39

114 \

/

\

FIGURE 4-7 TWIN ENGINE AIRCRAFT CONFIGURATION 4.2.2 Single Engine Airplane

Figure 4-8 shows the single engine aircraft. Characteristics are:

1. Propeller Data:

Prop. diameter 2.134 m Prop. speed at take-off 2,400 rpm Tip speed at take-off 268 m/sec = .79 a Prop. speed at economy cruise 1,800 rpm Tip speed at economy cruise 201 m/sec Prop. ground clearance 356 mm

2. Sight Angles

The centerline angle over the nose for the single engine airplane, C, is 9 ° . This should correspond to actual pilot's viewing angle of about 12 ° . This is probably adequate, especially when compared to some of today's long nose single engine'aircraft.

3. Aircraft Data

Single Engine Single Engine Diesel Gasoline Airframe minus engine (a) kg 667 667 Engine -- Figure 3-58 (b) kg 162 175 Empty weight (a) + (b) (c) kg 829 842 Payload (d) kg 340 333 Fuel load (e) kg 180 174 Useful load (d) + (e) (f) kg 520 507 Max. take-off weight (c) + (f) kg 1,349 1,349 Wing span m 11.16 11.16 Length m 8.66 8.66 Tail height m 3.14 3.14 Tail span m 3.78 3.78 Wing area m 2 17.74 17.74

116 C

-- C":

FIGURE 4-8 SINGLE ENGINE AIRCRAFT CONFIGURATION m ......

4.3 Aircraft Performance Evaluation

The major tool used in the airplane design synthesis was a somewhat modified version of the synthesis method originally developed for the NASA GATE (General Aviation Turbine Engine) Study. (13)* The process was simplified for this purpose since take-off and cruise power could be specified as program inputs. The program is not accurate enough nor does it account for enough variables to actually design airplanes, but it is considered adequate to indicate trends in relative size and performance for airplanes theoretically equipped with different engines. The main point to bear in mind when looking at the results of the program is that the objective is to provide an indication of the differences in performance and cost between diesel and gasoline powered airplanes. The methods used in estimating throughout are no better than 5 to 10% accurate, but the uniform assumptions and methods used in all cases would make the resulting differences good indications of the trends to be expected. This is the proper objective for a conceptual investigation.

4.3.1 Program Input Data

The data needed by th_ program can be put in three broad classifications:

1. Desired Airplane Mission Profile:

A. Payload.

B. Range and speed at cruise altitude.

C. Take-off and landing distances.

Mission Profile:

149 kW Single 298 kW Twin Cruise speed km/hr 324 474 Altitude m 3,048 7,620 Range km 1,481 2,592 Payload kg 340 726 Take-off distance m 579 701 Landing distance m 369 677 Cruise power kW 149 243 Take-off power kW 149 298

2. Engine Performance Data:

A. Take-off power.

B. Cruise power and fuel consumption at the specified cruise altitude.

C. Engine weight and geometry.

D. Propeller drive shaft speed for use in calculating propeller diameter and propulsive efficiency.

E. Induction airflows.

F. Cooling requirements.

118 Thefollowing tabulationgivesthe specific programinput data:

149 kW 298 kW Diesel Gasoline Diesel Gasoline BSFC g/kW-hr 228 268 213 286 Altitude m 3,048 3,048 7,620 7,620 Speed km/hr 315 315 444 444 Power % 100 75 81.5 75 Proposed rpm 2,400 2,300

The 298 kW develops full power up to 6,096 m. Above that altitude the power drops off in proportion to the ambient air density.

Engine weight data:

149 kW 298 kW Gasoline kg 175 262 Diesel kg 163 207 Difference kg 12 55

Cooling air estimates Engine cooling air requirements are used to calculate cooling air inlet areas, exit areas and momentum drag. Piston engine experience indicates that an v 1 inlet velocity ratio between the inlet and free stream of.4 is desirable (v0 - .4).

Similarly an exit velocity ratio of .3 is indicated (- Vex_ - .3). Vo

The inlet and exit areas are calculated using:

A - area m2 V A - where V = airflow volume m3/sec V

v = airflow velocity m/sec

Airflow volume V is determined from the weight WA:

WA = V - WA where weight airflow kg/sec d d= air density kg/m 3

119 Weight flow is determinedby the requiredheat rejection rates:

e heat rejection rate kcal/sec

WA _ Q where Cp -- spec. heat of air at constant pressure Cp T .24 kcal/kg/°C

AT= temp. rise of cooling air across heat exchanger 55.6°C

In calculating the cooling exit areas, a 4.4°C temperature rise was used in addition to the 55.6°C rise across the heat exchangers. This allowed for ram rise and radiation heating. Exit density used in calculating exit areas was determined by temperature ratio.

dex = exit densitykg/m 3

T Tex = exit temperature °K dex : d --where Tex d = density ambient air kg/m 3

T = temp. ambient air °K

The change in momentum of the air flowing through the heat exchangers in the engine compartment induces a drag force on the airplane. This is best represented in terms of thrust power required to provide the cooling air flow. It is a function of the velocity of the airplane and the atmospheric conditions:

TABLE XXIV Data for Cooling Air Duct Sizing and Cooling Air Momentum Drag Calculation

100% Power Cruise (Standard Ambient)

Engine Type Diesel Gasoline (1) Diesel (1) Gasoline Rated Power kW 149 149 298 298 Cruise Power kW 149 112 298 224 Speed km/h r 330 315 439 444 Altitude m 3,048 2,134 6,096 7,620 Cooling kcal/min Oil Cooler 252 363 403 255 After Cooler 390 -- 959 363 Cylinders 504 762 -- 762

120 Engine Type Diesel Gasoline (1) Diesel (1) Gasoline Fuel Injectors -- -- 151 -- Total Heat Rejection kcal/min 1,146 1,125 1,513 1,380 Induction Airflow kg/sec .32 .14 .56 .23 Inlet Area cm 2 432 -- 593 -- Exit Area cm 2 701 -- 980 --

(2) Drag (thrust kW) 10.2 8.2 21.8 20.9 (3) Cooling f m 2 .030 .025 .037 .040

(1) Each Engine (2) At altitudes and speeds listed. (3) Equivalent flat plate area for use at any altitude and speed.

Thrust power --_ WA Vo( Vo -Vex) kW 102

Input data and the results of the calculations are shown in Figure 4-8.

It should be noted that the cooling data for the gasoline engine refer to 75% cruise power while the diesel data apply to 100% cruise power. For a fair comparison the gasoline data should have been 33% higher.

3. Aerodynamic Characteristics and Weight Data:

These values are supplied by the program aerodynamicist using experience with the class of airplane being considered and the desired characteristics of the new design. These data include life and drag coefficients and the coefficients for an airplane weight calculation. Other values needed are tail size parameters, reserve fuel, air density, and constants used in take-off and landing distance calculations. Many of the constants and coefficients used are empirical. Some of the important values are shown in the following table:

TABLE XXV Aerodynamic Constants and Coefficients

(1) Single Engine Twin Engine Aircraft Aircraft C L Max. Landing 2.19 1.87

C L Max. Take-off 1.43 1.48

f Total Diesel (2) m 2 .30 .50 f Diesel-f-Gasoline Cooling: m 2 .0046 - .0033(5) Nacelle size m 2 _ .0084(5)

Reserve fuel(6) hours .75 .82

_CD/ACL2 (3) .0655 .0597

_P Cruise(4) .85 .85

121 (1)Thedataare for diesel andgasolineairplanesof constant size.The largerairplanesfor constant missioncomparisonsare scaledup as requiredfrom this basis. (2)Equivalentflat plateareausedto calculate profile drag. (3)Induceddragfactor. thrust power (4) Propeller Efficiency = shaft power (5) Each engine. (6) Includes allowance for climb, take-off and reserve.

In the case of the diesel twin, the drag was decreased by about 2% to allow for the smaller frontal and wetted area relative to the gasoline engine nacelle. The projected frontal area of the single engine airplane does not change since the cabin cross section stays the same.

The airplane size parameters obtained when the program is run using the above information include wing area, gross weight and fuel weight. Sets of data made up of inputs and resulting outputs allow synthesized airplanes of different sizes and with different engines to be" compared. The process is very simplified and is by no means a complete airplane design process but it does allow preliminary concepts to be evaluated side by side on the basis of the same set of assumptions.

4.3.2 Calculation Method

1. A trial airplane weight is selected.

2. Wing area required for landing is calculated using an empirical relation containing weight, wing lift and required landing distance.

3. Wing area required for take-off is calculated using an empirical relation containing weight, wing lift, power and required take-off distance.

4. Using the larger wing area from 2 or 3, cruise drag is calculated accounting for wing area, tail area, fuselage size, nacelle size and miscellaneous items.

5. Cruise power required is calculated to meet the speed requirement.

6. Fuel required to meet the range is then calculated.

7. Airplane weight is then calculated using an empirical relation accounting for fuel weight, payload, wing area and power.

8. The weight calculated in Item 7 is compared with the trial weight of Item 1. If different, a new trial weight is selected and the process repeated.

Using the data and methods described, hypothetical gasoline and diesel powered airplanes were synthesized and compared in two ways. In one case, the airframe was held constant and the mission profile was allowed to change when the power plant type changed. In the other case, the mission requirements were held constant _nd the airplane needed to perform that mission change d size as necessary to meet the mission requirements. These comparisons were made for both the single 149 kW engine and twin 298 kW engine airplanes.

122 4.3.3 Results of the Simulation Program

The results of the aircraft performance simulation program are shown in the Tables XXVl and XXVlI.

Table XXVI shows the differences in aircraft performance for a fixed airplane size.

The fixed parameters are: • Max. take-off weight • Max. landing weight • Take-off distance • Landing distance • Stall speed • Wing area

The advantages of the diesels with their high cruise power output and low fuel consumption can be readily seen in the basic parameters of range, speed, and payload.

TABLE XXVI Comparison Gasoline and Diesel Aircraft Engines Airplane Size Fixed, Variable Performance

Single-Engine Single-Engine Twin-Engine Twin Engine Diesel* Gasoline* Diesel* Gasoline* Rated power kW/RPM 149/2400 149/2600 298/2300 (ea) 298/2267 (ea) Max. take-off weight (gross) kg 1349 1349 3654 3654 Max. landing weight kg 1349 1349 3654 3654 Standard empty weight kg 829 842 2275 2385 Useful load kg 520 508 1378 1269 Usable fuel f./kg 251/180 241/174 908/653 832/598 Payload (with full fuel) kg 340 334 726 671

Altitude--m/%power 30481100% 3048/75% 7620181.5% 7620175% Max. cruise speed km/hr 324 291 474 448 Range km 1481 1468 2592 1726

Altitude -- m/% power 3048175% 3048/75% 7620/81.5% 7620/75% Speed km/hr 289 291 474 448 Range km 1968 1468 2592 1726

Take-off distance (normal, OV. 15 m) m 579 579 701 701 Landing distance (normal, OV. 15 m) m 369 369 677 677 Stall speed (landing) km/hr 85 85 135 135 Wing area m 2 17.7 17.7 22.4 22.4

*All engines are turbocharged.

123 TableXXVIIshowsthe differencesin airplanesize for a fixed performance. Thefixed parametersare: • Payload • Max.CruiseSpeed • Range

Thegasolinepoweredairplanesare bigger andconsiderablyless efficient.

TABLE XXVII Comparison Gasoline and Diesel Aircraft Engines Performance Fixed, Variable Airplane Size

Single-Engine Single-Engine Twin-Engine Twin-Engine Diesel* Gpsoline* Diesel" Gasoline* Rated power kW 149 198 298 (ea) 414 (ea) Max take-off weight (gross) kg 1349 1525 3654 4981 Max. landing weight kg 1349 1525 3654 4981 Standard empty weight kg 829 973 2275 3140 Useful load kg 520 552 1378 1842 Usable fuel ,_/kg 251/180 294/211 9081653 1552/1116 Payload (with full fuel) kg 340 340 726 726

Altitude- m/% power 3048/100% 3048/75% 7620/81.5% 7620/75% Max. cruise speed km/hr 324 324 474 474 Range km 1481 1481 2592 2592

Altitude- m/% power 3048175% 3048/100% 7620/81.5% 7620/75% Speed km/hr 289 324 474 474 Range km 1968 1481 2592 2592

Take-off distance (normal, OV. 15 m) m 579 564 701 701 Landing distance (normal, OV. 15 m) m 369 427 677 689 Stall speed (landing) km/hr 85 93 135 135 Wing area m 2 17.7 17.0 22.4 29.9

*All engines are turbocharged.

4.4 Operating Cost Estimates

Production costs were estimated by assuming that new airplanes would be designed and equipped with the diesel engines and, alternatively, compatible gasoline engines. Development, material and labor costs were chosen to be of roughly the correct magnitude, but are intended primarily to illustrate cost differences due to using diesel instead of gasoline engines. Operating cost estimates were made using figures obtained from current estimates of average operating costs.

4.4.1 Airplane Acquisition Cost Estimates

The acquisition cost estimates were based on information from the airplane synthesis process. The airplane empty weights were the main parameters used with FY79 rates

124 for labor, material costs and OEM engine costs. The estimating methods used are based on historical data and "learning curve" theory. An airframe weight was estimated from the operating empty weight. This was used with estimating data to get material weights to which material cost could be applied. Manhour per pound data were used to get labor content to which labor rates were applied. A production run of 6000 units was used to amortize assumed development costs and to locate factors on the learning curves. When a basic factory cost was summed up, assumed manufacturer's and dealer's mark-ups were applied. Costs were included for currently typical optional equipment and avionics selections. The final total represented a dealer's price tag figure for a typically equipped airplane. Both the single and the twin were considered to be all new designs. The same sets of reasonably realistic assumptions were used throughout so the results are quite adequate for looking at differences between gasoline and diesel airplane prices within the overall accuracy of this study. Acquisition price percentage changes from the diesel to the gasoline engine powered airplanes is shown on the cost summaries. See Tables XXVIII and XXX for the single and twin engine airplanes, respectively.

The factors used in calculating these costs are summarized in the Tables XXIX and XXXI.

TABLE XXVIII Cost Summary Single Engine

Use 500 Hours/Year

Equal Plane Performance Gasoline* Airplane Diesel Gasoline Acquisition cost Base -4% +11%

Fuel $/hr 8.54 9.90 13.20 Oil $/hr .43 .38 .51

Inspection & maintenance Airframe $/hr 2.59 2.59 2.59 Engine $/hr 4.00 2.59 2.59 Propeller $ / hr .30 .30 .30

Engine exchange $/hr 4.28 7.57 10.09 Hangar rental $/hr 2.40 2.40 2.40 Insurance $/hr 5.90 5.90 6.54

Total DOC/Hr. $/hr 28.44 31.63 38.22 Total per year $ 14220 15815 19110 Total for 5 years $ 71100 79075 95550

* Bigger airplane required to do the same job as the diesel.

125 TABLE XXIX Main Operating Cost Factors Summary Single Engine

Equal Plane Performance Factor Diesel Gasoline Gasoline Cruise speed @ 3048 m km/hr 324 291 324 Total cruise power output kW 149 112 149 BS FC g/kW/h r 228 268 268 Fuel density kg/,e .81 * .72 .72 Fuel cost $/,_ .20* .24 .24 Oil density kg/,_ .87 .87 .87 Oil cost $/_** 1.11 1.11 1.11 Engine exchange cost $ 12841 10604 14103t hours 3000 1400 1400

*. **Oil consumption is 1% of fuel consumption. t$/Rated kWratio( 1§8 )from 149 kW gasoline engine. 149

TABLE XXX Cost Summary Twin Engine Use 1000 Hours/Year

Equal Plane Performance Airplane Diesel Gasoline Gasoline* Acquisition cost Base - 3% +7%

Fuel $/hr 26.79 42.30 58.69 Oil $/hr 6.98 1.53 2.13

Inspection & maintenance Airframe $/hr 9.20 9.20 9.20 Engine $/hr 13.80 13.80 13.80 Propellers $/hr 2.00 2.00 2.00

Engine exchange $/hr 19.82 34.84 48.34 Hangarrental $/hr 3.30 3.30 3.30 Insurance $/hr 6.18 5.99 6.59

Total DOC/Hr. $/hr 88.07 112.96 144.05 Total per year $ 88070 112960 144050 Total for 5 years $ 440350 564800 720250

Bigger airplane required to do the same job as the diesel.

126 TABLE XXXI Main Operating Cost Factors Summary Twin Engine

EqualPlane Performance Factor Diesel Gasoline Gasoline Cruise speed @ 7620 m km/hr 474 448 474 Total cruise horsepower kW 501 447 620 BS FC g/kW/h r 213 286 286 Fuel density kg/,_ .81 * .72 .72 Fuel cost $/_ .20* .24 .24 Oil density kg/_. .94* * .87 .87 Oil cost $/,_t 6.34* * 1.11 1.11 Engine exchange cost $ 24775 20907 29008tt Time between overhauls hours 2500 1200 1200

*Jet fuel. * *Synthetic oil. tOil consumption is 1% of fuel consumption. tt$/Rated kW ratio ( 414 ) from 298 kW gasoline engine. 298

The columns headed "gasoline" refer to the airplanes of equivalent size to the diesels but with these mission capability as indicated in the performance estimates. The "equal plane performance gasoline" column refers to the airplanes that will do the same missions as the diesels but are bigger and less efficient.

The cost summary pages show the considerable overall cost advantage of the diesel powered airplanes. Gasoline airplanes of equivalent size cost less initially but this advantage is not significant in view of the reduced mission capability and higher overall costs. The biggest factors in raising the gasoline airplanes operating costs are fuel and overhaul expense, as indicated.

4.5 Propeller Noise Estimates

Propeller performance estimates were made to get some idea of the propeller sizes needed to realize a cruise propulsive efficiency of .85 for both the single and twin engine airplanes. These calculations indicated that a two-blade, 2134mm diameter, constant speed propeller will work for the single engine airplane. The propellers indicated for the twin are 2057mm three-blade. Estimates of 305m flyover noise predict values of 72 dB(A) for the single and 74 dB(A) for the twin. These compare favorably to the limits of 77.5 riB(A) and 80 dB(A), respectively. Limits are based on airplane weight as set out in FAR 36, Appendix F. A favorable correction factor can reasonably be expected, creating a greater margin relative to the limits. The correction factor is based on detailed take-off performance estimates that are beyond the scope of this study. Even without correction factors, the noise regulations appear to present no problem for the conceptual diesel airplanes.

127 5.0 CONCLUSIONS

The study indicates that the diesel engine promises to be a superior powerplant for general aviation aircraft.

1. The diesel engine offers high cruise power at altitude and low fuel consumption. This will result in improved range, high cruising speed and more payload for a diesel engined aircraft.

2. The diesel powered airplane has a considerable overall cost advantage. Gasoline airplanes of equivalent size cost less initially, but this advantage is offset by reduced mission capability and higher operating costs.

3. The diesel engine presents no installation problems. Although the radial configuration is different than current gasoline engines, the mounting to the airframe is essentially the same and requires no major airframe modifications.

4. The engine can run on diesel fuel and jet fuel.

5. The independent turbo loop provides these features:

A. Easy cold and hot starts.

B. Can crank engine indefinitely.

C. Electric power available independent of engine operation (APU mode).

D. Reduced battery capacity.

E. Cabin cooling or heating available while aircraft is on the ground.

6. The radial cylinder configuration results in:

A. Low engine weight.

B. Reduced engine friction.

C. Absence of piston inertia forces.

D. Compactness of the power package.

7. The two-stroke cycle feature results in:

A. Weight reduction.

B. Improved reliability due to fewer parts.

C. Reduced frontal area.

128 8. Alternatesolutions areavailablefor the high risk technologies:

A. Limited cylinder cooling can be substituted for uncooledcylinderswhich requirethe use of ceramiccomponents.

B. A conventionalcombustorcan be substituted for the catalytic combustor.

C. An enginedrivenalternatorcan replacethe high speedturbo driven alternator.

129 6.0 RECOMMENDATIONS

The program has shown the feasibility of the diesel engine as a powerplant for general aviation aircraft, the technologies which were applied to the engine designs are currently under development under various Government contracts but require more experience and adaptation to an aircraft engine. It is recommended that development programs be initiated starting with single cylinder test engines and leading to full scale multi-cylinder engines for test cell performance and testing and eventual flight experience.

130 7.0 LIST OF REFERENCES

1. P.H. Wilkinson, "The Performance of Modern Aircraft Diesels," SAE Transactions, Vol. 47, No. 5.

2. H. Sammons and E. Chatterson, "The Napier Nomad Aircraft Diesel Engine," SAE #320, 1954.

. J. Dooley, "McCulloch is Developing Lightweight Aircraft Diesel," SAE Transactions, Vol. 79, No. 9.

4. R. Blaser, A. Pouring, E. Keating, B. Rankin, "The Naval Academy Heat Balanced Engine," Report USNA EW#8-76.

° J.R. Grundy, L.R. Kiley, and E.A. Brevick, "A VCR-1360-2 High Specific Output Variable Compression Ratio Diesel Engine," SAE Paper #760051.

6. D.F. Mowbray and M. Drori, "The CA V DP15 Fuel Injection Pump," SAE Paper #780163.

7. J.H. Stang, M.E. Woods, W.C. Geary, A.S. Williamson and W.A. Updike, "Development of an Adiabatic Diesel Engine," USATARADCOM, Report #12345.

. M. Berchtold and F.J. Gardiner, "The Comprex, A New Concept of Diesel Supercharging," ASME Paper #58-GTP-16.

9. J. Melchior and T. Andre-Talamon, "Hyperbar System of High Supercharging," SAE Paper #740723.

10. P.S. Patel and E.F. Doyle, "Compounding the Truck Diesel Engine with an Organic Rankine Cycle System," SAE Paper #760343.

11. P.H. Schweitzer, "Scavenging of Two-Stroke Cycle Diesel Engines." Publisher: The MacMillan Company, New York.

12. H. List, "Der Ladungswechsel der Verbrennungskraftmaschine." Publisher: Springer-Verlag, Vienna.

13. W.A. Strack, "New Opportunities for Future Small Civil Turbine Engines -- Overviewing the GATE Studies," SAE Paper #790619.

131 APPENDIX A

Bibliography

° J.F. Alcock, J.V.B. Robson and C. Mash, "Distribution of Heat Flow in High Duty Internal Combustion Engine," CIMAC, 1957.

. M. Alperstein, G.H. Schofer and F.J. Villforth, "Texaco Stratisfied Charge Engine Multi-Fuel, Efficient, Clean and Practical, "SAE Paper 740563.

, F.L. Arnold and J.S. Prestley, "Hypereutectic AI-Si Casting Alloys Phosphorus Refinement," Modern Castings, March 1961.

4. C.F. Bachle, "Progress in Light Aircraft Engines,'.' SAE Transactions, Vol. 46, 1940.

° J.M. Bailey, "Multifuel Combustion System for High Performance Diesel Engines," SAE Paper 790 B, 1964.

° A.J.S. Baker, D. Dowson and P. Sterachan, "Dynamic Operating Factors in Piston Rings," JSME Conference, Tokyo, Nov. 1973.

7. I. Balint and L. Brinson, "Two-Stage Turbocharging and Intercooling," ASM E Paper 68 -- DG P-5, 1968.

8. H.W. Barnes-Moss, "Engine Design for the Future," SAE Paper 741130.

9. H.W. Barnes-Moss and WoM. Scott, "The High Speed Diesel Engine for Passenger Cars," SAE Paper 750331.

10. S.G. Berenyi, "Variable Area Turbocharger Design, A VCR 1360 Engine," TCM/GPD Report, Contract DAAE 07-73-C-0292, April 1974.

11. R. Bertode, T.W.E. Downes and I.D. Middlemiss, "Evaluation of a New Combustion System for Diesel Emission Control, " SAE Paper 741131.

12. I.N. Bishop, "The Effect of Design Variables on Friction and Economy, "SAE Transactions, Vol. 73, 1965.

13. J.W. Bjerklie, E.J. Cairns, C.W. Tobias and D.G. Wilson, "An Evaluation of Alternative Power Sources for Low Emission Automobiles," SAE Paper 750929.

132 14.G.P.Blairand W.L.Cahoon, "Designand Initial Development of a High Specific Output 500 cc Single Cylinder Two-Stroke Racing Motorcycle Engine," SAE Paper 710082.

15. V. Bock and A. Ostergaard, "The Development of a High Pressure Charged 4-Stroke Engine," CIMAC, 1965.

16. B. Brisson, E. Ecomard and P. Eyzat, "A New Diesel Combustion Chamber -- The Variable Throat Chamber," SAE Paper 730167.

17. E.M.W. Britain, "Thermal Loading of Small High-Speed Two-Stroke Diesel Engines," Symposium Inst. Mech. Engrs., Birmingham, 1964, Paper #7.

18. D. Broome, "Toward Higher Speeds and Outputs from the Small Diesel Engine," SAE Paper 730149.

19. D.H. Brown, "Diesel Cylinder Heat Transfer Design Criteria," SAE Paper 249, 1957.

20. W. Bryzik, "Adiabatic Diesel Engine," U.S. Army Tank Automotive R & D Command.

21. J.F. Cassidy, "Electric Close-Loop Controls for the Automobile," SAE Paper 740014.

22. E. Chatterton, "The Napier Deltic Diesel Engine," SAE Paper 632, 1955.

23. D.L. Clason and W.H. Nahumck, "Performance Evaluation of a Versatile Two-Cycle Lubricant Additive, "SAE Paper 790080.

24. R.W. Claus, J.D. Wear and C.H. Liebert, "Ceramic Coating Effect on Liner Metal Temperatures of Film-Cooled Annular Combustor," NASA Technical Paper 1323.

25. H.R.M. Craig and M.S. Janota, "The Potential of Turbochargers as Applied to Highly Rated Two-Stroke and Four- Stroke Ehgines," CIMAC, 1965.

26. K. Dao, O.A. Uyehara and P.S. Myers, "Heat Transfer Rates at Gas Wall Interfaces in Motored Piston Engine," SAE Paper 730632.

27. P.K. Das, "Analysis of Piston Ring Lubrication," SAE Paper 760008.

133 28. J.G. Dawson, W.J. Hayward and P.W. Glamann, "Some Experiences with a Differentially Supercharged Diesel Engine," SAE Paper 932 A, 1964.

29. R.A. DeKeyser, M.A. Gates and J.H. Parks, "Caterpillar's New Sleeve Metering Fuel Injection Systems," SAE Paper 750079.

30. J. Dickson and R.D. Wellington, "General Motors 'Model 51' Diesel Engine," SAE Paper #27, 1953.

31. J. Dickson and R.D. Wellington, "Loop Scavenging Used in New 3000 RPM Diesel," SAE Journal, Vol. 61, 1953.

2, J.L. Dilworth, "The Sky is the Limit for Two-Cycle Engines," SAE Paper 599 A, 1962.

33. P.K. Doerfler, "Comprex Supercharging of Vehicle Diesel Engines," SAE Paper 750335.

34. C.T. Doman, "Economic Aspects of Light Airplane Engines," SAE Transactions, Vol. 47, #1.

35. J.L. Dooley and J.W. Wittlesey, "A Highly Compounded 2-Cycle Radial Light Plane Engine," SAE Paper 660173.

36. J.L. Dooley, "Two-Stroke Light Aircraft Engine Potential," SAE Paper 670238.

37. G.Eichelberg and W. Pflaum, "Untersuchung eines Hochaufgeladenen Dieselmotors," VDI, Vol. 93, 1951.

38. G. Eichelberg and W. Pflaum, "A New MAN Engine," ASNE Journal, 1953.

39. E. Engdahl, "The Application of Chemical Vapor Deposited Silicon Carbide to Radial Turbomachinery," SAE Paper 760282.

40. E. English, "Piston Rings for High Duty Internal Combustion Engines," CIMAC, 1965.

41. Ethyl Corp., "Aviation Fuels and Their Effect on Engine Performance," NAVAER 06-5-501, USAF TQ # 06-5-4.

42. D.B. Field and S.J. Hinkle, "Detroit Diesel Allison's Series 92 Engines," SAE Paper 740037.

43. C.C.J. French, E.R. Harles, and M.L. Monaghan, "Thermal Loading of Highly Rated Two-Cycle Marine Diesel Engines," CIMAC,1965.

134 44. C.C.J.French, "Taking the Heat Off the Highly Boosted Diesel," SAE Paper 690463.

45. C.C. French and D.H. Taylor, "An Investigation into Diesel Engine Operation at Very High Ratings (500 psi BMEP)," CIMAC, 1975.

46. S.L. Gaal, J.P. Peer, G.L. Muntean and H.L. Wilson, "A New Injector Concept for the D.I. Diesel to Improve Smoke and Fuel Consumption at Low Emission Levels," CIMAC, 1975.

47. G. Garratt and D.E. Gee, "The Application of Some Analytical Techniques to the Study of Turbocharging the Automotive Diesel Engine," Proc. Inst. Mech. Engrs., 1965-1966, Vol. 180, Pt. 3 N.

48. E.J. Gay, "Powerplants for Industrial and Commercial Vehicles _ A Look at Tomorrow," SAE Paper 650477.

49. G.E. Gazza and R.N. Katz, "Development of Advanced Sinterable Si3 N_," DOE Meeting, Oct. 1978, Dearborn, Michigan.

50. I. Gorille, N. Rittmannsberger and P. Werner, "Bosch Electronic Fuel Injection with Closed-Loop Control," SAE Paper 750368.

51. H.H. Haas and E.R. Klinge, "The Continental 750 HP Aircooled Diesel Engine," SAE Transactions, Vol. 65, 1957.

52. H.E. Helms and F.A. Rockwood, "Ceramic Applications in Turbine Engines," DOE Meeting, Oct. 1978, Dearborn, Michigan.

53. N.A. Henein, "Instantaneous Heat Transfer and Coefficients Between Gas and Combustion Chamber of a Diesel Engine," SAE Paper 969 B, 1965.

54. N.A. Henein, "The Diesel as an Alternative Automobile Engine," SAE Paper 750931.

55. Th. P. Herbell and.Th. K. Glasgow, "Developing Improved Reaction Sintered "Silicon Nitride," DOE Meeting, Oct. 1978, Dearborn, Michigan.

55. S.S. Hetrick, J.A. Keller and H.V. Lowther, "Performance Advantages of Synthesized Commercial Engine Oils," SAE Paper 780183.

57. W. Hinze, "Zur Wirkungsgradminderung durch den Waermeverlust bei Dieselmotoren," MTZ, Jan. 26, 1965.

135 58. S.D. Hines and G.L. Pochmara, "An Analytical Study of Exhaust Gas Heat Loss in a Piston Engine Exhaust Port," SAE Paper 760767.

59. E.W. Hives, and F.L. Smith, "High Output Aircraft Engines," SAE Transactions, Vol. 46, #3, 1939.

60. H.G. Holler, "Tomorrow's Diesel _ What Will It Offer," SAE Paper 650479.

61. R.J. Hooker, "Orion _ A Gas Generator Turbocompound Engine," SAE Paper 794, 1956.

62. J.l. Hope and R.D. Johnston, "A New Concept for Reduced Fuel Consumption in Internal Combustion Engines," SAE Paper 719051.

63. J.H. Horlock and R.S. Benson, "The Matcffing of Two-Stroke Engines and Turbochargers," CIMAC, 1962.

64. H.U. Howe and F. Pischinger "Der Luftgekuehlte Deutz Diesel Motor FL 912," MTZ #29, 1968.

65. W.J. Hull, "High Output Diesel Engines," SAE Transactions, Vol. 72, 1964.

66. D.W. Hutchinson, "The Differential Gas Turbine," SAE Transactions, 1956.

67. G.F. Hyde, J.E. Cromwell and W.C. Arnold, "Piston Ring Coating for High Performance Diesel Engines," SAE Paper 670935.

68. M.S. Janota, "An Experimental Survey of Component Temperatures in a Small High Speed Loop Scavenged Two-Stroke Diesel Engine with Varying Simulated Turbocharging Conditions," Symposium Inst. of Mech. Engrs., Birmingham, Oct. 1964, Paper #9.

69. D.J. Jordan and H.S. Crim, "Evolution of Modern Air-Transport Powerplants," Journal of Aircraft Sept./Oct. 1964.

70. J.L. Jorstad, "The Hypereutectic AI-Si Alloys 390 and A390," AIME Paper.

71. R. Kamo and W. Bryzik, "Adiabatic Turbocompound Engine Performance Prediction," SAE Paper 780068.

72. R. Kamo, "Preliminary Report on Vehicle Testing of Cummins Turbocompound Diesel Engine," DOE Presentation, Oct. 1978, Dearborn, Michigan.

136 73. R.N. Katz and E.M. Lenoe, "Ceramics Technology," ERDA Meeting, Oct. 1976, Ann Arbor, Michigan.

74. E.S. Kopecki, "Tool Steels for Automotive Engines and Other Components," SAE Paper 780247.

75. E.H. Kraft, "Silicon Carbides for DOENASA Sponsored Automotive and Industrial Gas Turbine Program," DOE Presentation, May 1978, Troy, Michigan.

76. E.M. Lenoe, "Current Status of Life Prediction Methodology for Ceramics," DOE Meeting, May 1978, Troy, Michigan.

77. C.H. Liebert, "Emittance and Absorptance of NASA Ceramic Thermal Barrier Coating System," NASA Technical Paper 1190.

78. H. List, "High Speed, High Output, Loop Scavenged Two-Cycle Diesel Engines," SAE Paper #855, 1956.

79. W.G. Lundquist, "Aircraft Powerplants _ Present and Future," SAE Paper #486, 1955.

80. H.H. Macklin, "New Developments for Aluminum Engines," 4th National Die Casting Congress, Nov. 1966

81. W.P. Mansfield, "A New Servo-Operated Fuel Injection System for Diesel Engines," SAE Paper 650432.

82. W.P. Mansfield, D.W. Tryhorn and C.H. Thornycroft, "Development of the Turbocharged Diesel Engine to High Mean Effective Pressure Without High Mechanical or Thermal Loading," CIMAC, 1965.

83. W.P. Mansfield and W.S. May, "Diesel Combustion at High MEP with Low Compression Ratio, " SAE Paper 660343.

84. J. Marley, "Simplification of System Inputs and Outputs for MPU Control Units," SAE Paper 780123.

85. T. Matsumura, "Development of a Turbocharged Two-Cycle Aircooled Diesel Engine," SAE Paper 720783.

86. A.F. McLean and D.A. Davis, "The Ceramic Gas Turbine _ A Candidate Powerplant for the Middle and Long Term Future," SAE Paper 760239.

137 87. W.E. Meyer and J.J. DeCarolis, "Compression Temperatures in Diesel Engines Under Starting Conditions," SAE Transactions, Vol. 70, 1962.

88. S.J. Miley, E.J. Cross, N.A. Ghomi and P.D. Bridges, "Determination of Cooling Air Mass Flow for a Horizontally, Opposed Aircraft Engine Installation," SAE Paper 790609.

89. B.W. Millington and E.R. Hartles, "Frictional Losses in Diesel Engines," SAE Paper 680590.

90. J.E. Mitchell, "An Evaluation of Aftercooling in Turbocharged Diesel Engine Performance, " SAE Transactions, Vol. 67, 1959.

91. V. Montaner, A. Antonucci and P.F. Rivolo, "A New Diesel Injection Pump with High Injection Rate, Its Influence on Smoke and Emissions," SAE Paper 750744.

92. NMAB Committee on Structural Ceramics, "Structural Ceramics," National Materials Advisory Board Report NMAB-320, 1975.

93. H.V. Nutt, W.W. Landen and J.A. Edger, "Effect of Surface Temperatures on Wear of Diesel Engine Cylinders and Piston Rings," SAE Transactions, 1955.

94. The Oil Engine and Gas Turbine, "Altitude Tests of Turbocharged Engines," Oct. 1961.

95. R.T. Paluska, J.S. Saletzki and G.E. Cheklich, "Design and Development of a Very High Output (VHO) Multifuel Engine," SAE Paper 670520.

96. R.C. Pampreen, "The Use of Variable Inlet Guide Vanes for Automotive Gas Turbine Engine Augmentation and Load Control," SAE Paper 760285.

97. J. Panton, "The Radial Turbine as a Ceramic Component," ASME Paper 78-GT-178.

98. R.F. Parker, "Future Fuel Injection System Requirements of Diesel Engines for Mobile Power," SAE Paper 760125.

99. W.G. Payne and W.S. Lang, "High Supercharging Development of a GM16-278A Two-Stroke Cycle Diesel Engine," SAE Transactions, Vol. 63, 1955.

100. W.H. Percival, "Method of Scavenging Analysis for 2-Stroke Cycle Diesel Cylinder," SAE Transactions, 1955.

138 101.H.F.Prasse,H.E.McCormickandR.D.Anderson, "AutomotivePiston Rings _ 1967 State of the Art," SAE Paper 670019.

102. Private Pilot, 1967, "What's Ahead in Lightplane Powerplants? - - - And Now, Two Strokes."

103, J.M. Radford, W.B. Wallace and R.A. Dennis, "Experimental Techniques Used in the Development of Highly Rated 4-Stroke Cycle Diesel Engines," CIMAC, 1965.

104. C.J. Rahnke, "The Variable Geometry Power Turbine," SAE Transactions, Vol. 78, 1969.

105. Ricardo & Co., Ltd., "Questions Raised by the 520 Aircraft Diesel," NASA/U of M Project.

106. R.W. Richardson, "Automobile Engines for the 1980's," Eaton's Worldwide Analysis of Future Automotive Powe'rplants, 1973.

107. J.G. Rivard, "Closed Loop Electronic Fuel Injection Control of the Internal Combustion Engine," SAE Paper 730005.

108. W.Rizk, "Experimental Studies of the Mixing Processes and Flow Configurations in Two- Cycle Engine Scavenging," Proc. Inst. Mech. Engrs., 1958.

109. R.R. Robinson and J.E. Mitchell, "Development of a 300 psi BMEP Continuous Duty Diesel Engine," CIMAC, 1965.

110. M.D. Roehrle, "Pistons for High Output Diesel Engines," SAE Paper 770031.

111. A.R. Rogowski and C.L. Bouchard, "Scavenging a Piston -- Ported 2-Stroke Cylinder," NACA Tech. Note #674, 1938.

112. V.D. Roosa, T.D. Hess and J.W. Walker, "The Roosamaster Nozzle," SAE Paper 907 B, 1964.

113. J.H. Rush, "Exhaust Port Heat Rejection in a Piston Engine -- A Preliminary Report," SAE Paper 760766.

114. R.B. Schultz, "Ceramics in the ERDA Highway Vehicle Heat Engine Systems Program," SAE Paper 760238.

115. W.J. Schultz, C.E. Miesiak, A.E. Hamilton and D.E. Larkinson, "Credibility of Diesel over Gasoline Fuel Economy Claims by Association," SAE Paper 760047.

139 116. P.H. Schweitzer, "Research in Exhaust Manifolds," CIMAC, 1951.

117. P.H. Schweitzer and L.J. Grunder, "Hybrid Engines," SAE Transactions, Vol. 71, 1963.

118. P.H. Schweitzer, A.W. Hussmann and A.E. Sieminski, "Lycoming S & H. The Compact Multifuel Engine," SAE Paper 790 C, 1964.

119. K.W. Self, "The Future of Higher Horsepower Engines," SAE Paper 384 A, 1961.

120. G.M. Shulhan and H.I.H. Saravanamuttoo, "Variable Geometry Compressors for Improvement of Gas Turbine Part Load Performance," SAE Paper 760283.

121. R.E. Siewert and S.R. Turns, "The Staged Combustion Compound Engine (SCCE): Exhaust Emissions and Fuel Economy Potential," SAE Paper 750889.

122. E.W. Spannhake, "Procedures Used in Development of Barnes & Reinecke Air Force Diesel Engine," SAE Transactions, Vol. 61, 1953.

123. J.H. Stang, "Designing Adiabatic Engine Components," SAE Paper 780069.

124. R.R. Stephenson, "Should We Have a New Engine," JPL SP 43-17, Vol. II, 1975.

125. A. Stotter, K.S. Woolley and E.S. Ip, "Exhaust Valve Temperature. A Theoretical and Experimental Investigation," SAE Paper 969 A, 1965.

126. C.F. Taylor and A.R. Rogowski, "Scavenging the 2-Stroke Engine," SAE Transactions, Vol. 62, 1954.

127. C.F. Taylor, A.R. Rogowski, A.L. Hagen and J.D. Koppernaes, "Loop Scavenging vs. Through Scavenging of 2-Cycle Engines," SAE Paper #247, 1957.

128. C.F. Taylor and T.Y. Toong, "Heat Transfer in Internal Combustion Engines," ASME Paper 57-HT-17, 1957.

129. D.W. Tryhorn, "Turbocharging the Automotive Two-Stroke Cycle Engine," Proc. Inst. Mech. Engrs., 1965-66.

130. K.C. Tsao, P.S. Myers and O.A. Uyehara, "Gas Temperatures During Compression in Motored and Fired Diesel Engines," SAE Transactions, Vol. 70, 1962.

140 131. O.A. Uyehara et al, "A Classification of Reciprocating Engine Combustion Systems," SAE Paper 741156.

132. D.R. Vance and R. Mayernick, "lntercooled vs. Non-lntercooled Engine Performance," SAE Paper 766 A, 1963.

133. W.T. Van der Nuell, "Turbocharging for Better Vehicle Engines," SAE Paper 631 A, 1963.

134. P.S. Vaughan, "The Development of a High Specific Output 4-Stroke Supercharged Diesel Engine," CIMAC, 1965.

135. E.T. Vincent and N.A. Henein, "Thermal Loading of Wall Temperatures as Functions of Performance of Turbocharged Compression Ignition Engines," SAE Transactions, Vol. 67, 1959.

136. E.T. Vincent, "Performance Prediction for the Highly Turbocharged Compression Ignition Engine," SAE Paper 766 B, 1963.

137. C.J. Walder, "Some Problems Encountered in the Design and Development of High Speed Diesel Engines," SAE Paper 978 A, 1965.

138. F.J. Wallace, "The Differential Compound Engine," SAE Paper 670110.

139. F.J. Wallace and D.E. Winterbone, "Performance Prediction of Very High Output Two and Four Stroke Engines Using a Generalized Computer Program," Proc. Inst. Mech. Engrs., 1970.

140. F.J. Wallace, P.C. Few and P.R. Cave, "The Differential Compound Engine _ Further Development," SAE Paper 710085.

141. F.J. Wallace and P.R. Cave, "Experimental and Analytical Scavenging Studies on a Two-Cycle Opposed Piston Diesel Engine," SAE Paper 710175.

142. F.J. Wiegand and M.R. Rowe, "Compounding the Piston Engine," SAE Paper 335, 1949.

143. F.J. Wiegand and W.R. Eichberg, "Development of the Turbocompound Engine," SAE Transactions, Vol. 62, 1954.

144. C.H. Wiegman, "Geared Engines for Light Airplanes," SAE Transactions, Vol. 47, 1940.

145. W.A. Wisemann and E.J. Ounsted, "TIARA Light Aircraft Engines _ A New Generation," SAE Paper 700205.

141 146. J.E. Witzky, "Forecasts of Future Powerplant Developments," ASME Paper 64-OGP-8, 1964.

147. G. Woschni, "Heat Transfer in Internal Combustion Engines," MTZ 4/26/1965.

148. F.V. Zalar and D.E. Nisbett, "Current Developments in Diesel Engine Oil Technology," SAE Paper 780182.

149. K.D. Zimmermann, "New Robert Bosch Developments for Diesel Fuel Injection," SAE Paper 760127.

150. K. Zinner, "Theoretical and Experimental Investigation of an Operational Procedure Involving the Use of a Coupled Exhaust Turbine," CIMAC, 1962.

142 APPENDIX B Metric Conversion Factors From: Multiply by: To:

kW 1.341 HP

mm .0394 inch

61.024 in3 cu.litersdm t

kg 2.2046 Ib

km .6214 mile

kPa .145 psi

m/sec 196.85 fpm

kW/kg .6083 HP/Ib

kW/cm 2 8.656 HP/in 2

kW/liter .022 HP/in 3

km .5401 nautical mile

g/kW-hr .00164 Ib/HP-hr

kg/I 62.453 Ib/fP

kcal 3.9683 BTU

N-m .7375 ft-lb

MPa 145 psi

kcal/min-kW 2,959 BTU/min-HP

m 3.2808 ft

liter ,264 gallon

kcal/kg 1.8 BTU/Ib

kW 56.826 BTU/min

lq3 3. Recipient's Catalog No. 1. Report No. / 2. Government Accession No. NASA CR-3260 1 5. Report Date 4. Title and Subtitle 150AND300kwLIG. VE OHT iESELAIRC FTENGINE April 1980 6. Performing Organization Code DESIGN STUDY

Performing Organization Report No. Author(s) 756 Alex P. Brouwers 10. Work Unit No.

9. Performing Organization Name and Address

Teledyne Continental Motors 11. Contract or Grant No. 76 Getty Street NAS3-20830

Muskegon, Michigan 49442 13. Type of Report and Period Covered

12. Sponsoring Agency Name and Address Contractor Report

National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, D.C. 20546

15. Supplementary Notes NASA Lewis Final report. Project Manager, Lloyd W. Ream, Engine Systems Division, Research Center, Cleveland, Ohio 44135.

16. Abstract This design study reintroduces the diesel engine as an aircraft powerplant. A methodical study was conducted to arrive at engine configurations and applicable advanced technologies. Two engines are discussed, a 300 kW six-cylinder engine for twin engine general aviation aircraft and a 150 kW four-cylinder engine for single engine aircraft. The description of each engine includes concept drawings, a performance analysis, stress and weight data, and a cost study. This information was used to develop two airplane concepts, a six-place twin and a four-place single engine aircraft. The aircraft study consisted of installation drawings, computer gener- ated performance data, aircraft operating costs, and drawings of the resulting airplanes. The performance data show a vast improvement over current gasoline-powered aircraft. A second report, NASA CR-3261, covers a design, performance, and cost study of a 186 kW aircraft diesel engine applicable to single and twin engine aircraft. A 5 year program consisting of component development and single-cylinder and multicylinder performance and endurance tests of the 186 kW engine is covered in CR-3261.

18. Distribution Statement 17. Key Words (Suggested by Author(s)) I Unclassified - unlimited Aircraft diesel engine; Diesel aircraft engine; Ad- I STAR Category 07 vanced engines; Adiabatic diesel engine; Diesel with I independent turbocharger loop; High speed starter/

alternator for aircraft diesel; Two-stroke cycle air- I craft diesel; Radial diesel aircraft engine J 22. Price* 19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 147 A07 Unclassified Unclassified

* For sale by the NationalTechnical InformationService, Springfield, Virginia 22161 NASA-Langley, ]9_